The Aviation Maintenance Alerts provides a common communication channel through which the aviation community can economically interchange service experience, cooperating in the improvement of aeronautical product durability, reliability, and safety. This publication is prepared from information submitted by those who operate and maintain civil aeronautical products. The contents include items that have been reported as significant, but have not been evaluated fully by the time the material went to press. As additional facts such as cause and corrective action are identified, the data will be published in subsequent issues of the Alerts. This procedure gives Alerts’ readers prompt notice of conditions reported via a Malfunction or Defect Report (M or D) or a Service Difficulty Report (SDR). Your comments and suggestions for improvement are always welcome. Send to: FAA; ATTN: Aviation Data Systems Branch (AFS-620); P.O. Box 25082; Oklahoma City, OK 73125-5029.

(Editor’s notes are provided for editorial clarification and enhancement within an article. They will always be recognized as italicized words bordered by parentheses.)

AIRPLANES

BEECH

Beech: A200; Twisted Cabin Door Bellcrank; ATA 5210

A technician for an air taxi operation states, "During a routine, Phase 3 inspection the main cabin door aft up-lock bell crank was found twisted about 10 degrees from (that) of a new bell crank (P/N 50-430031-17). The up-lock bracket pivot bolt (P/N NAS1103-21D) was also noted to be slightly deformed. This (condition) was probably caused by the up-lock hook being (improperly) rigged. All effected parts were replaced and the up-lock hooks were (then) able to be rigged properly."

(Slightly deformed? Nice comparison photos.)

Part Total Time: (unknown).

Beech: B65-A80; Sheared Nose Gear Actuator; ATA 3220

"The nose landing gear would not deploy on approach to landing," says an air taxi submitter. "After nose gear removal it was discovered that the nose gear actuator drive shaft (P/N 50-820218) internal spline was worn and sheared. (Lubrication maintenance should be consistent...) to reduce wear."

Part Total Time: 4,383.0 hours.

Beech: B200C; Cracked Wing Skin; ATA 5730

A mechanic states, "During a pre-flight inspection, maintenance technicians noted a 3.75 inch long crack in the right wing lower wing skin, just forward of the inboard section of the outboard flap. Additionally, a support/anchor tab that attaches the lower wing skin to the inboard most flap-well rib of the outboard wing ...was found cracked through (failed). We believe this tab failure allowed the skin to vibrate, generating rapid crack growth. This crack is (positioned) directly aft of the lower aft wing bolt. (It) had just appeared within the past two hours of flight time. I can pin-point this time frame since the 5 year wing bolt removal and inspection process was completed just 2 flight hours ago. No wing skin cracks were noted or documented at that time." (R/H trailing edge lower wing skin P/N: 50-110027-15)

Part Total Time: 8,914.0 hours.

CESSNA

Cessna: 208; Chafed Engine Mount; ATA 7120

(This aircraft is being pulled by a Pratt & Whitney PT6A-114A.)

A mechanic writes, "During (this) aircraft's corrosion inspections, it was noticed that the air conditioning condenser duct clamp was cutting into the engine mount on the L/H truss. The truss was also rubbing against the condenser duct. There are four cuts measuring as follows (length x depth; respectively): 9/32 inch x .020—2 each; 7/32 inch x .010; 9/16 inch x .032. I had determined the clamp was correctly installed, and that the part numbers are in accordance with the current Cessna 208 parts catalog. The clamp, when installed on the engine mount, rubs against the truss and creates damage. Cessna was notified by e-mail and (they) recommend using aluminum 'speed tape' under the clamp to reduce damage to the truss. Cessna has not provided an official engineering analysis, and the recommendation was provided by e-mail. This engine mount assembly was replaced. We would like to see this clamp replaced by a more suitable clamp to prevent damage to the engine mount truss. (Operators of...) the Cessna Caravan fleet should be notified of the possible unsafe condition created by the rubbing/cutting of the clamp, possibly causing failure of the engine mount."

(Engine mount P/N: 2651023-19. A search of the FAA Service Difficulty Reporting System (SDRS) database reflects a second, very similar defect report. Readers should note part total time!)

Part Total Time: 575.2 hours.

Cessna: 208B; Burned Flap Motor Wires, ATA 2752

A mechanic writes, " We departed (our base) at approximately 8:30 local time. The flaps were set at 10 degrees.... Take-off and climb were normal. After about 10 to 15 minutes into the flight (with the autopilot engaged) we noticed a burning smell and saw a small amount of smoke—lasting for about a second or two. We disengaged the autopilot and looked at the circuit breakers, noticing the stall warning circuit breaker had popped. After returning to (base) we found several circuit breakers had popped and that the flaps were inoperative, stuck at 10 degrees. An avionics shop was contacted to look at the aircraft—they reported back the following (account)."

"A hidden damage inspection was performed. Burned wires were found in the wire bundle running from the main flap motor to the circuit breaker panel and from the main flap motor to the ground block J66. Insulation of adjacent wires along the same wire bundle were also found melted. A loose ring terminal was found on the normally open terminal of the up limit switch on the flap actuator assembly. The flap system was inspected for any defects structurally—none were found. The flap actuator jackscrew exhibited no abnormal signs of wear—it rotated freely when checked with the primary motor removed. Upon further investigation, it was found that the primary flap motor had failed, and the 10 amp primary flap circuit breaker had not tripped as it should have, causing the circuit to overload and melt the wire bundle. The primary flap circuit breaker exhibited signs of overheating on the bottom surface of the plastic casing, and had a burned odor to it. The primary flap motor looked good physically, but upon further investigation (opening the motor) it exhibited serious signs of overheating/burning inside the unit (this also having a very strong burned odor). During the investigation to determine the severity of the damage caused by the failed circuit, it is estimated at least 70 to 75 percent of the wire insulation on the power wire to the primary flap motor had been melted/burned off of the wire, leaving a bare exposed wire throughout the cabin headliner. The wire finally burned itself in half behind the circuit breaker panel about two inches from the J37 connector. The following components were removed and sent to Cessna for investigation: Flap actuator (P/N 9910586-3); Flap Circuit breaker (P/N S1232-510); Flap Control Relays K20 and K21 (P/N MS24187D1); and the Flap overhead Normal/Standby power switch (P/N MS25068-23). We are waiting to hear from Cessna as to what was the root cause of this (defect)."

"During the SID (shaft inspection device) inspection," says an unknown submitter, "the fuselage bulkhead at the forward opening of the cabin door and secondary bulkhead were found cracked, buckled, and distorted. (The only reason...) this was found was the SID inspection requires removal of the door bayonet receptacles for inspection, and (also) the removal of the interior sidewall upholstery."

(Bulkhead P/N 57111434345. Nice camera/mirror work—Ed.)

Part Total Time: 7,246.0 hours.

MITSUBISHI

Mitsubishi: MU2B25; Cracked Gearbox Housing; ATA 7210

A repair station technician writes, "A pilot reported low and fluctuating oil pressure from the left engine while in flight, and elected to shut the engine down. An uneventful single engine landing was made (at a landing site). The engine oil bypass valve was found to be in the 'popped' condition. The oil filter was then removed—(it was observed) to be black in color, with the forward sealing surface broken in several small pieces. The reduction gearbox was disassembled and the oil pump removed. The forward (reduction gearbox) housing (P/N 3102047-4) was cracked approximately three inches radially at the forward detail."

An unidentified submitter states, "These Pneumatic Control Timers (Lamar P/N 488-699) were purchased new...and when they were installed in the aircraft it was noted the surface deice boots were inflated (continuously) when the aircraft power was on. A second timer unit was purchased and the same problem existed when it was installed. Upon troubleshooting the system and verifying correct installation and wiring, it was found that two of the wires coming out of the unit were switched during manufacture and the boxes were operating opposite as is intended. There are four wires: red, black, white, and blue. The white and blue wires are reversed."

Part Total Time: 0.00 hours.

HELICOPTERS

HILLER

Hiller: UH-12E; Incorrect Transmission Retainer; ATA 6330

"(I) suspect this (defective retainer part) is locally made," says a repair station technician. The installation word 'up' is not stamped on this part (retainer; P/N 23688) but is etched on by hand. The thickness of this retainer is .045 inches—the actual part thickness of a new retainer is .051 inches. (I) found shafts that are shimmed with bad retainers to be shimmed improperly due to (this) retainer thickness—as is required by the Hiller overhaul manual. The shafts were out of tolerance by .014 to .020 inches." (Transmission P/N: 23700-23.)

A private pilot provides the following report. "During replacement of the canister-type oil filter at the time of oil change, it is tricky to achieve a good seal of the canister against the seat (with) the filter gasket that is supplied with the Champion CFO100-1 filter kit. This gasket very easily slips inside the canister during tightening of the canister retainer bolt. Extreme care must be taken to ensure it does not slip inside. If it does slip, there is a strong oil leak during the run-up check. This failure of proper gasket sealing has happened—not only to me as a (CFR) Part 91 private pilot changing my own oil, but also to both mechanics in the shop I use. So far, removal, cleaning, and drying of the gasket has resulted in a good seal during reinstallation. A better design could save us lost oil, and having to pay the mechanic for extra time and clean-up. This has been an ongoing issue for at least three years."

(Good evaluation! The SDRS program can use all the input from private pilots it can get—Ed.)

Part Total Time: (n/a).

AIR NOTES

INTERNET SERVICE DIFFICULTY REPORTING (iSDR) WEB SITE

The Federal Aviation Administration (FAA) Internet Service Difficulty Reporting (iSDR) web site is the front-end for the Service Difficulty Reporting System (SDRS) database that is maintained by the Aviation Data Systems Branch, AFS-620, in Oklahoma City, Oklahoma. The iSDR web site supports the Flight Standards Service (AFS), Service Difficulty Program by providing the aviation community with a voluntary and electronic means to conveniently submit in-service reports of failures, malfunctions, or defects on aeronautical products. The objective of the Service Difficulty Program is to achieve prompt correction of conditions adversely affecting continued airworthiness of aeronautical products. To accomplish this, Malfunction or Defect Reports (M or Ds) or Service Difficulty Reports (SDRs) as they are commonly called, are collected, converted into a common SDR format, stored, and made available to the appropriate segments of the FAA, the aviation community, and the general public for review and analysis. SDR data is accessible through the “Query SDR data” feature on the iSDR web site at: http://av-info.faa.gov/sdrx/Query.aspx.

In the past, the last two pages of the Alerts contained a paper copy of FAA Form 8010-4, Malfunction or Defect Report. To meet the requirements of *Section 508, this form will no longer be published in the Alerts; however, the form is available on the Internet at: http://forms.faa.gov/forms/faa8010‑4.pdf. You can still download and complete the form as you have in the past.

*Section 508 was enacted to eliminate barriers in information technology, to make available new opportunities for people with disabilities, and to encourage development of technologies that will help achieve these goals.

A report should be filed whenever a system, component, or part of an aircraft, powerplant, propeller, or appliance fails to function in a normal or usual manner. In addition, if a system, component, or part of an aircraft, powerplant, propeller, or appliance has a flaw or imperfection, which impairs or may impair its future function, it is considered defective and should be reported under the Service Difficulty Program.

The collection, collation, analysis of data, and the rapid dissemination of mechanical discrepancies, alerts, and trend information to the appropriate segments of the FAA and the aviation community provides an effective and economical method of ensuring future aviation safety.

The FAA analyzes SDR data for safety implications and reviews the data to identify possible trends that may not be apparent regionally or to individual operators. As a result, the FAA may disseminate safety information to a particular section of the aviation community. The FAA also may adopt new regulations or issue airworthiness directives (ADs) to address a specific problem.

The iSDR web site provides an electronic means for the general aviation community to voluntarily submit reports, and may serve as an alternative means for operators and air agencies to comply with the reporting requirements of 14 Title of the Code of Federal Regulations (CFR) Section 121.703, 125.409, 135.415, and 145.221, if accepted by their certificate-holding district office. FAA Aviation Safety Inspectors may also report service difficulty information when they conduct routine aircraft maintenance surveillance as well as accident and incident investigations.

The SDRS database contains records dating back to 1974. At the current time, we are receiving approximately 40,000 records per year. Reports may be submitted to the iSDR web site on active data entry form or submitted hardcopy to the address below.

You can access current and back issues of this publication from the internet at: http://av‑info.faa.gov/. Select the General Aviation Airworthiness Alerts heading.

AVIATION SERVICE DIFFICULTY REPORTS

The following are abbreviated reports processed for the previous month, which have been entered into the FAA Service Difficulty Reporting (SDR) System database. This is not an all-inclusive listing of Service Difficulty Reports. For more information, contact the FAA, Regulatory Support Division, Aviation Data Systems Branch, AFS-620, located in Oklahoma City, Oklahoma. The mailing address is:

To retrieve the complete report, click on the Control Number located in each report. These reports contain raw data that has not been edited. Also, because these reports contain raw data, the pages containing the raw data are not numbered.

If you require further detail please contact AFS-620 at the address above.

Federal Aviation Administration

Service Difficulty Report Data

Sorted by aircraft make and model then engine make and model. This report derives from unverified information submitted by the aviation community without FAA review for accuracy.

DURING ENGINE CHANGE THE OVERSPEED GOVERNOR WAS REMOVED AND CHIPPED GEAR TOOTH WAS DISCOVERED. STATUS SHEET INDICATED THE GOVERNOR WAS INSTALLED 2006/05/26 AND HAS ACCUMULATED 2312.2 TIME SINCE OVERHAUL. THE ENGINE, THE GOVERNOR WAS INSTALLED ON, HAS BEEN SENT FOR OVERHAUL AS SCHEDULED AND THE OVERSPEED GOVERNOR HAS BEEN SENT FOR REPAIR.

TIRE WAS FOUND LOW ON PRESSURE UPON DAILY INSPECTION. ATTEMPT TO INFLATE TO PROPER PRESSURE WAS DONE, BUT TIRE FAILED TO MAINTAIN PRESSURE. WHEEL ASSEMBLY REMOVED FROM SERVICE. FURTHER INSPECTION REVEALED THAT TIRE HAS PRESSURE LEAKS FROM AROUND BEADSEAT AREA ON INNER AND OUTER SIDES OF TIRE. ALSO DISCOVERED ON INSPECTION THAT LEAKS WERE DISCOVERED FURTHER UP THE SIDEWALL TO TREAD. ALSO THAT THIS TIRE IS A RECAPPED TIRE IN ITS 1ST SCHEDULED RECAP.

IT WAS NOTICED ON A RETURN FLIGHT THAT OUR AIRCRAFT HAD OIL STAINING AND LEAKAGE ON THE LOWER LEFT NACELLE AND LANDING GEAR. WHEN THE AIRCRAFT WAS TROUBLESHOT IT WAS FOUND THAT THE STEWART WARNER FUEL TO OIL HEATER WAS CRACKED ON THE CORNER EDGE RADIUS FOR ROUGHLY AN INCH IN LENGTH. IF YOU OBSERVE THE FOLLOWING PICTURES YOU CAN SEE THE BUBBLING FROM THE LEAK WHEN THE PART WAS PRESSURIZED.

THE MAGNETO WAS REMOVED FOR A SCHEDULED MAINTENANCE CHECK ON THE DISTRIBUTOR BLOCK AND BRUSH ASSEMBLY. THE IMPULSE COUPLING STOP PIN WAS OBSERVED WORN BEYOND MANUFACTURERS LIMITS OF 0.050 INCH. MEASURED WEAR WAS 0.155 INCH MAGNETO HAS LESS THAN 30 HOURS SINCE NEW.

ACFT ON GROUND RUNNING APU FOR ABOUT 3 HOURS BEFORE DEPARTURE. ON TAKEOFF, SMALL AMOUNT OF SMOKE IN CABIN AND COCKPIT. RETURNED TO AIRPORT AND LANDED. PERFORMED AN INSP OF APU AREA AND FOUND SMALL AMOUNT OF OIL IN AREA. PERFORMED GROUND POWER ENGINE RUNS AND NO SMOKE NOTED. PERFORMED OPS FLIGHT TO CHECK SYS. SYSTEM CHECKED OK. NO SMOKE NOTED. (K)

FOUND THESE FITTINGS AN-916 1D TO BE DEFECTIVE. THE FITTINGS WERE BEING USED ON A HYDR SYS WITH A LOW PRESSURE OF 1,100 PSI. WHEN ASSEMBLED ORIGINALLY THE FITTINGS WERE ASSEMBLED WITH A LIGHT COAT OF FUEL LUBE. WHEN FOUND TO LEAK, THE FITTING ASSY WAS REMOVED, TIGHTENED AND REASSEMBLED. ONCE AGAIN, WE HAD A LEAK. STATED TIGHTENING EACH FITTING OF A RUGN AND THEN PLACED ON A HAND PUMP HYD MULE. EACH TIME IT LEAKED. HAD (1) SPARE SET OF FITTINGS, MADE UP THE SAME ASSY AGAIN AND THIS TIME BEGAN TO GO PAST WHAT WAS CONSIDERED A REASONABLE AMOUNT OF TORQUE FOR THE ASSY. USING THE SAME PROCEDURE OF TIGHTENING THE FITTING A TURN AND PUTTING IT BACK ON THE MULE. EACH TIME FLUID STARTED COMING OUT AROUND BOTH SIDES OF THE FITTING AT LESS THAN 200 PSI. BOTH ASSYS WERE MADE USING THE AN-9161D WITH A AN8164D ON ONE SIDE AND AN AN8254D ON THE OTHER. THESE WERE INSTALLED ON THE PLANE WHEN LEAKS WERE DISCOVERED. THE OTHER ASSY WAS MADE UP OF THE SPARE THAT WAS ORDERED USING THE SAME PARTS AS THE FIRST ASSY. ALL 3 FITTING LEAK AROUND BOTH SIDES OF THE AN916-1D. THE LAST FITTING ASSEMBLED WAS OVER TIGHTENED TO SEE IF LEAK WOULD STOP. RAN OUT OF MALE THREADS BEFORE LEAK WOULD STOP. SEEMS LIKE THE PIPE THREAD TAP WAS RUN TOO FAR INTO THE FITTING DURING THE MFG PROCESS OR THERE IS NOT ENOUGH TAPER ON THE THREAD. HAVE LABELED THE FITTINGS 1-3. NR 1 IS THE FITTING ORIGINALLY INSTALLED ON THE PLANE WITH THE AN 816-4D. NR 2 IS THE ONE THAT WAS INSTALLED ON THE PLANE IN A LOWER PRESSURE AREA. NR 3 IS THE ONE ASSEMBLED AFTER AND KEPT TIGHTENING AND TESTING. NR 3 IS THE ONE THAT THE MALE THREADS WERE COMPLETELY RUN INTO ON BOTH ENDS. WAS UNABLE TO REMOVE THE AN-816-4D FROM THE NR 3. DISTRIBUTOR WHO SUPPLIED PARTS INFORMED MECH THAT HAD RECALLED PARTS FROM LOT NR THAT WAS TRYING TO BE PUT ON ACFT. LOT NR ON BAG. (K)

AT THE END OF THE 3 HOUR BENCH RUN FROM OVERHAUL AN INTERNAL INSP OF THE MAGNETO WAS ACCOMPLISHED. THE TIP OF THE TACHOMETER POINTS (10-400507) WAS FOUND IN THE COMPARTMENT. THE CONTACT BROKE OFF THE RIVET ON THE BREAKERS SPRING SIDE. THE CONTACTS WERE NEW WITH DATE STAMP TCM-0750 AND THE ONLY RUN TIME WAS ON THE BENCH. INSPECTION UNDER MAGNIFICATION WOULD INDICATE A JOINT FAILURE OF THE WELD BETWEEN THE RIVET AND SILVER CONTACT..

HAVE FOUND (5) NEW 4596 LANDING LIGHT BULBS WITH CORRODED WIRE ATTACHMENT TERMINALS. CORROSION COLORING WAS EITHER WHITE OR GREEN, FLAKING TYPE POWDER. ALL BULBS WERE NEW IN BOX. TOO MUCH SOLDIER FLUX?

DATA PLATE INDICATES D-3000 MAGNETO, HOWEVER, MAGNETO HSG HAS BEEN POSITIVELY IDENTIFIED TO BE D-2000. HSG ALSO MODIFIED FOR D-3000 DISTRIBUTOR BLOCK. AD APPLYING TO D-2000 MAGNETO HSG ARE ELIMINATED WITH D-3000 HSG. USING THE D-2000 HSG IS POTENTIALLY DANGEROUS. AD APPLICABLE TO D-2000 HSG: 78-18-06, 80-17-14. THESE AD IF NOT COMPLIED WITH BECAUSE OF MISREPRESENTATION OF DATA PLATE COULD RESULT IN FAILURE OF MAGNETO. (K)

WE RECEIVED NEW DISTRIBUTOR BLOCK AND GEAR KIT FOR BENDIX MAG S6RN-1225.UPON ASSEMBLING THE GEAR AND DISTRIBUTOR BLOCK THE HIGH TENSION TERMINAL OF THE GEAR WAS FOUND TO BE RUBBING THE DISTRIBUTOR BLOCK AT THE TIMING MARK LOCATION ON THE BLOCK. THIS HAPPENED ON 4 OF THE 5 SETS OF PARTS WE RECEIVED. PARTS WERE TAGGED AND RETURNED TO VENDOR.

ENGINE, SN: LE-45724AEF (REFERENCE WO NR 240967, SVS NR 45724E) WAS RETURNED TO THIS REPAIR STATION. WHILE ENGINE WAS OPERATIONAL THERE WAS A LOUD BANG JUST BEFORE THE ENGINE SHUT DOWN IN FLIGHT. THE AIRFRAME WAS A DUAL ENGINE DESIGN AND PILOT MANAGED TO PERFORM AN UNSCHEDULED LANDING (OEI) WITHOUT INJURIES OR DAMAGE TO THE ACFT. UPON DISASSEMBLY OF THE ENGINE AND SEPARATING THE POWER MODULES WE DISCOVERED THE GAS PRODUCER (GP) NOZZLE (PN 4-111-090-02, SN 972025600047 TSN 6650 HOURS) TO HAVE SEPARATED RADIALLY (ABOUT 270 DEGREES) ON THE CURL PORTION OF THE NOZZLE ASSEMBLY. WE ARE NOT CONCLUDING THIS AS THE CAUSE OF THE INCIDENT. WE ARE JUST REPORTING A PART CONDITION WE HAVE NOT WITNESSED BEFORE. UPON NOTIFYING THE ENGINE DESIGN HOLDER. WE WERE ASKED TO STOP ALL WORK AND TO SEND THE POWER MODULES TO THEM. WE PACKAGED AND SHIPPED THE POWER MODULES FOR FURTHER DISASSEMBLY AND INVESTIGATION.

DURING HOT SECTION INSP, 2 COMPRESSOR TURBINE BLADES, PN 3045741-01, WERE FOUND TO HAVE CRACKS IN THE T/E OF THE BLADE. THESE CRACKS WERE FOUND ABOUT HALF WAY UP THE T/E. ONE CRACK .3750 INCH IN LENGTH. THE OTHER WAS .1875 INCH IN LENGTH. CRACKS WERE FOUND BY FLUORESCENT PENETRANT INSPECTION. SB1703 WAS GOING TO BE COMPLIED WITH, BUT AFTER CRACKS WERE FOUND, ALL BLADES WERE SCRAPPED. COMPRESSOR TURBINE BLADES TIME SINCE NEW: 6940.1 TIME SINCE OVERHAUL: 1782.9

(CAN) SHORTLY AFTER T/O, THE ENGINE SUFFERED AN UNCOMMANDED DECREASE IN POWER. THE FLIGHT WAS ABORTED AND THE AIRCRAFT RETURNED TO BASE WITHOUT INCIDENT. THE DECREASE IN POWER WAS COMPENSATED BY THE OTHER ENGINE AND AT NO TIME DID THE ENGINE REVERT TO MANUAL NOR WAS THERE ANY FAULT INDICATED ON THE COCKPIT DISPLAY. TROUBLESHOOTING LED TO THE REPLACEMENT OF THE ENGINE HARNESS DUE TO AN ERRATIC TORQUE SIGNAL. POST REPLACEMENT GROUND CHECKS AND TEST FLIGHT WERE ALL NORMAL. MFG WILL CONTINUE INVESTIGATING THE EVENT AND ADVISE OF ROOT CAUSE ONCE ESTABLISHED.

SUBJECT ENGINE WAS RECEIVED DUE TO HIGH TIME. DURING DISASSEMBLY AND INSP OF THE HPC THE DIFFUSER AIRSEAL PN 56H192, SN CENCBG1041 WAS FOUND WITH LINEAR CRACKS IN THE BOLT HOLE FLANGE ID, IAW CIR MANUAL PN 51A750 INSP CHECK 02 (NDI) NO CRACKS ARE ALLOWED. ENGINEERING INVESTIGATING. (K)

AFTER TAKEOFF AT 200 FT AGL, THE RT ENGINE LOST OIL PRESSURE. THE CREW SHUT DOWN THE ENGINE AND DECLARED AN EMERGENCY AND RETURNED TO THE AIRPORT. UPON INVESTIGATION, MAINTENANCE FOUND THE WHEELCASE SCAVENGE FILTER LYING IN THE BOTTOM OF THE COWL. THE FILTER WAS COMPLETE WITH THE SINGLE ATTACHMENT BOLT. ATTACHED TO THE BOLT WAS THE INSERT THAT CONTAINED THE THREADS TO WHICH THE BOLT THREADS. IN SHORT, THE INSERT PULLED OUT OF THE CASING CAUSING THE FILTER TO COME OUT AND DEPLETING ALL ENGINE OIL. THE INCIDENT WILL BE PUT THROUGH THE COMPANY SMS SYSTEM FOR ACTION.

RT LANDING GEAR WOULD NOT LOCK DOWN. PILOT RETURNED TO DEPARTURE WHILE TROUBLESHOOTING ISSUE AND EXECUTING EMERGENCY CHECKLIST. ADDITIONAL TROUBLESHOOTING OVER RADIO WITH PILOT AFTER CONSULTING WITH MX. GEAR WOULD NOT LOCK IN FULL DOWN AND LOCK POSITION. ELECTED TO HAVE SOMEONE TRY TO PULL THE GEAR DOWN WHILE THE PILOT HOVERED. DID EXTENSIVE COORDINATION PRIOR TO THE ATTEMPT AND WAS SUCCESSFUL. GEAR LOCKED DOWN AND ACFT WAS TOWED TO THE HANGAR.

WE HAVE BEEN EXPERIENCING A VIBRATION DURING DESCENT. AFTER SUBSTANTIAL INSPECTIONS, TESTING AND CONSULTATION WITH ATR, A CHECK WAS CARRIED OUT TO MEASURE THE SPRING TAB FOR WARPAGE. THE RESULTS WERE SENT TO ATR FOR ANALYSIS. THE WARPAGE LIMITS WERE DEEMED OUT OF LIMITS. THE SPRING TAB WAS REPLACED AND THE VIBRATION CEASED DURING SUBSEQUENT FLIGHTS.

ON TWO SEPARATE OCCASIONS WHILE PERFORMING UNSCHEDULED MAINTENANCE INVOLVING THE MLG SELECTOR VALVE , TWO ALUMINUM HYDRAULIC FITTINGS HAVE BEEN FOUND DAMAGED OR CRACKED (1) OCT 27, 2008 - WHILE TROUBLESHOOTING A REPORTED HYDRAULIC LEAK UNION P/N MS21902D6 HAD A DAMAGED FLARE. (2) JAN 9, 2009, WHILE REPLACING THE MLG SELECTOR VALVE REDUCER P/N ER804D060604 WAS FOUND CRACKED. SUSPECT THAT THESE FITTING HAVE BEEN OVER TORQUED CAUSING DAMAGE AND STRESS CRACKS.

DURING ENGINE RUN-UP AFTER C CHECK CREW HEARD LOUD BANG FROM LT ENGINE. THIS OCCURRED BETWEEN 70 PERCENT AND 60 PERCENT TQ, DURING PL REDUCTION FROM 80 PERCENT TQ TO GI. AFTER SHUT DOWN, VISUAL INSPECTION REVEALED HEAVY DAMAGE ON THE 2ND STAGE POWER TURBINE, IMPACT DAMAGE ON THE EXHAUST PIPE AND PROPELLER COULD NOT BE ROTATED MANUALLY. METAL DEBRIS WAS FOUND APPROX 20M BEHIND THE ENGINE. THE ENGINE WILL BE BOROSCOPE TO ACCESS INTERNAL DISTRESS. P&WC WILL CONTINUE INVESTIGATING THE EVENT AND ADVISE OF ROOT CAUSE ONCE ESTABLISHED.

(CAN) DURING CRUISE, THE CREW NOTED MAIN OIL PRESSURE FLUCTUATION WHICH WAS SHORTLY FOLLOWED BY A LOUD NOISE AND A LOSS OF POWER. THE ENGINE WAS SHUTDOWN AND THE FLIGHT WAS DIVERTED, WHERE A SINGLE ENGINE LANDING FOLLOWED. POST FLIGHT INSP, FOUND OIL COVERING COWLINGS AND MAJOR METALLIC CONTAMINATION OF THE TURBOMACHINE CHIP DETECTOR. THE ENGINE WILL BE REPLACED. MFG WILL CONTINUE INVESTIGATING THE EVENT AND ADVISE OF ROOT CAUSE ONCE ESTABLISHED.

(CAN) WHILE TAXING ONTO GATE, MX NOTICED FUEL COMING OUT FROM THE COWLING. CLOSER INSP REVEALED THE OUTER SLEEVE OF THE FLEXIBLE HOSE TO THE LOW PRESSURE PUMP WAS TORN DUE TO AN INTERNAL LEAK BETWEEN THE INNER LINER AND THE STAINLESS STEEL BRAIDING. THIS WAS A PRE SB RA320735 HOSE ASSY. THE HOSE WAS REPLACED.

AT THE TRANSITION FROM DECENT TO LEVEL FLIGHT ON A SEEDING PASS, PROP RPM WAS INCREASED AND A VIBRATION WAS FELT. AS THE THROTTLE WAS ADVANCED THE VIBRATION INCREASED DRAMATICALLY. THE ACFT WAS FORCED TO LAND AND AT LOW SPEED, FLIPPED OVER AS IT BROKE THROUGH A FOOT OF SNOW. ON SCENE INVESTIGATION REVEALED THAT ONE OF THE COUNTERWEIGHT BEARING SHAFTS HAD SHEARED, AND WAS NO LONGER PENETRATING THE PROPELLER BLADE.

DURING ANNUAL INSP, NUMEROUS CRACKED L/E RIBS WERE FOUND BETWEEN 66.3750 INCH TO 173.7500 INCH ON THE LT WING AND 86.6250 INCH TO 201 INCHES ON THE RT WING. THESE CRACKS WERE IN THE RIB BODY AND RAN ALONG THE UPPER SKIN ATTACH FLANGE. THE CRACKS STARTED AT THE AFT END OF THE FLANGE. LENGTHS WERE BETWEEN .0625 INCH TO .5 INCH. CORRECTIVE ACTION WILL CONSIST OF INSTALLING P.N 21062-(1 OR 2) AND P.N. 21068-(1 OR 2) DOUBLERS IAW DRAWING 21065 REPAIR - LE RIB - 802 AFTER THE INSTALLATION OF ADDITIONAL INSPECTION HOLES IAW DRAWING 21066 INSTALLATION - ADDITIONAL INSP. PLATE - WING L/E - 802.

ON ACFT, ACCOMPLISHING SB 7X-106 DRAFT 1 TITLED "ELECTRICAL POWER-PRIMARY DC GENERATION SYS INSP OF THE RT MAIN FEEDER GENERATOR 2 INCH. THE BLACK BOOT WHICH SURROUNDS THE FWD FEEDER CABLE FEED THROUGH CONNECTION AT FR46LH SIDE WAS FOUND SLIGHTLY MELTED ON THE TOP SIDE. (PG8 FG 2 OF SB) THE CUSTOMER OF THE ACFT THEN WANTED A FULL INVESTIGATION OF FEEDER CONNECTION AT FR 46 WITH THE ENGINE REMOVED. AFTER THE ENGINE REMOVAL THE MAIN WIRE FEED-THROUGH CONNECTOR AT FR 46 WAS DISSEMBLED. ARCING AND OVERHEATING WAS FOUND UNDER THE FWD LUG CONNECTION AND FEED THROUGH ASSEMBLY. ALL FINDINGS WERE REPORTED TO MFG ENGINEERING. REPAIR/ REPLACEMENT OF THE NR 2 POWER FEEDER CABLE FROM FR41 POWER DISTRIBUTION BOX TO THE FR 46 BULKHEAD CONNECTOR WITH A NEW DESIGNED FR 46 BULKHEAD CONNECTOR WAS ACCOMPLISHED IAW MFG TECH INSTRUCTIONS TI-7X-R0032-M. ALL OF THE ABOVE ACCOMPLISHED AT SERVICE CENTER. DURING AN 1A INSP.

DURING ANNUAL INSPECTION TO TWO 4-MAN LIFE RAFTS, BOTH LIFE RAFTS WERE FOUND TO HAVE THEIR COMPRESSED GAS CYLINDERS DEPLETED OF ALL STORED GAS PRESSURE. NEITHER RAFT WOULD HAVE WORKED IF NEEDED IN AN EMERGENCY. BOTH RAFTS WERE STILL COMPLETELY PACKED AND APPEARED TO BE UNDISTURBED AFTER THEIR LAST MX. (MAY 2008).

UPON REMOVAL OF THE HORIZONTAL STABILIZER, IT WAS NOTED THE FUSELAGE ATTACH FITTING CHANNEL APPARENTLY HAD BECOME FILLED WITH WATER, FROZE AND BURST. A (3) INCH CRACK ACROSS THE SPAN OF THE ATTACH FITTING SHOWED SIGNS OF BULGING OUTWARD. INSPECTION IN THIS AREA IS LIMITED WITHOUT STABILIZER REMOVAL.

UPON REMOVAL OF VERTICAL FIN FROM HORIZONTAL STABILIZER FOR ROUTINE MX, THE VERTICAL FIN ATTACH BRACKET IN THE STABILIZER WAS FOUND LOOSE AND HAD ELONGATED MOST ALL OF THE MOUNTING HOLES IN THE STABILIZER SPAR. WITH THE TAIL FULLY ASSEMBLED, NO LOOSENESS HAD BEEN NOTED. THIS AREA HAD NEVER BEEN ACCESSED, THE LOOSENESS CAME FROM IMPROPER TORQUE FROM THE FACTORY AT ASSY.

(CAN) UPON DESLUGING THE INNER BORE OF THE CRANKSHAFT, IT WAS DISCOVERED THAT THE PID COATING HAD BEGUN TO BUBBLE AND CHIP OFF. THE PID WAS REMOVED AND SURFACE CORROSION WAS DISCOVED UNDER THE PID SIMMILAR TO THAT REFERENCED IN AD 980208 AND SB 505B. IT SHOULD BE NOTED THAT THERE IS ONLY ONE CALANDER YEAR ON THIS ENGINE SINCE OVERHAUL.

UPON REMOVAL OF THE EXPANSION PLUG TO FACILITATE SLUG REMOVAL IT WAS DISCOVERED THAT THEIR WAS BUBBLING IN THE PID. BASED UPON PREVIOUS EXPERIENCE WITH THIS PROBLEM ON OTHER AIRCRAFT THE PID COATING WAS REMOVED AND PITTING SIMILAR TO THAT FOUND IN SERVICE BULLETIN NR 505B WAS FOUND.

AIRCRAFT RETURNED TO BASE NOTING HIGH FUEL CONSUMPTION ON RT ENGINE. AFTER LANDING LARGE AMOUNT OF FUEL NOTED LEAKING FROM RT NACELLE. FUEL LINE WAS RUPTURED FROM FCU TO START CONTROL. REF SDR 20080513005, 20080211006 AND 20080521010. THIS WAS THE FOURTH LINE FAILURE. AIRCRAFT RETURNED TO SERVICE WITH A LINE REMOVED FROM A SPARE ENGINE. AFTER INVESTIGATION FOUND NR 2 BEARING TO HAVE EXCESSIVE MOVEMENT AND NR 2 BEARING SEAL LEAKING. ENGINE WAS REMOVED AND REPLACED WITH OVERHAULED ENGINE. IT IS THOUGHT THE VIBRATION FROM THE NR 2 BEARING CAUSED A VIBRATION THAT RESONATED THROUGH THE LINE CAUSING THE LINE TO FAIL.

THE TREND MONITORING PROGRAM STARTED TO SHOW SIGNS OF HOT SECTION DISTRESS ON THE GRAPHS. BOROSCOPE INSPECTION REVEALED SOME EROSION AND IMPACT DAMAGE ON THE CT BLADES. AN HOT SECTION INSPECTION WAS CARRIED OUT AND THE COMPONENTS REPLACED. WHEN THE CT BLADES AND STATOR WAS INSPECTED AT THE APPROVED OVERHAUL FACILITY IT WAS DETERMINED THAT THE DAMAGE WAS CONSISTENT WITH EXCESSIVE CARBON ERODING THE TURBINE WITH IMPACT DAMAGE FROM LARGER PIECES OF CARBON WHICH DETACHED FROM THE COMBUSTION LINER. THE COMBUSTION LINER HAD ALSO BEEN REPLACED AT THE HOT SECTION INSPECTION DUE TO EROSION. THE MANUFACTURER HAS NOT SEEN THIS TYPE OF DAMAGE BEFORE ON THIS TYPE OF ENGINE.

(CAN) DURING 6TH, 200HR INSP, NOTED STRINGER P/N 114-440000-33 CRACKED ON AFT END APPROX. 3 INCHES LONG. THIS STRINGER IS LOCATED IN THE AFT FUSELAGE ASSY. AND SOME OF THE ANCHOR NUTS FOR THE LT STABILON ARE ATTACHED TO THIS STRINGER. THE STRINGER ON THE RT SIDE WAS INSPECTED AND NO FAULTS FOUND. STRINGER REPLACED WITH NEW.

(CAN) WHEN WE REPLACED THE HIGH PRESSURE COMPRESSOR BLEED VALVE FOR TOUBLESHOOTING, THE OVERHAULED UNIT THAT WAS INSTALLED CAUSED THE ENGINE TO HAVE SEVERAL COMPRESSOR STALLS AT IDLE. THE OVER HAULED BLEED VALVE WAS REMOVED FROM THE ENGINE AND EXAMINED. NOTICED THAT THE PISTON IN THE BLEED VALVE WAS SEIZED IN THE CLOSED POSITION, THE HSG WAS ASSEMBLED USING SILICONE SEALER AND SOME OF IT GOT ON TO THE PISTON. THE VALVE HAD BEEN OVERHAULED ON NOV 2005 AND HAD NEVER BEEN INSTALLED.

(CAN) REPLACED THE LOW PRESSURE COMPRESSOR BLEED VALVE FOR TROUBLE SHOOTING, THE OVERHAULED UNIT THAT WAS INSTALLED CAUSED THE ENGINE TO HAVE SEVERAL COMPRESSOR STALLS AT IDLE. THE OVERHAULED BLEED VALVE WAS REMOVED FROM THE ENGINE AND EXAMINED. NOTICED THAT PISTON IN THE BLEED VALVE SEIZED IN THE CLOSED POSITION, THE VALVE HSG WAS ASSEMBLED USING SILICONE SEALER AND THE SPRINGS WERE BROKEN INSIDE. THE VALVE HAD BEEN OVERHAULED ON NOV 2005 AND HAD NEVER BEEN INSTALLED.

THE BUSHING PN 101-521046-5 THAT HOLDS THE BEARING PN4A INSIDE THE RT AILERON BELLCRANK PN 101-521046-3 WAS FOUND LOOSE IN THE BELLCRANK AND INDUCED A LOSS MOTION OF APPROX .75 INCH AT THE T/E WITH THE RIG PINS INSTALLED.

THE SIDE PANEL LIGHTING WOULD NOT WORK UNTIL THE RHEOSTAT KNOB WAS PULLED. WHEN THE KNOB WAS WIGGLED A CRACKLING SOUND COULD BE HEARD. THE RHEOSTAT WAS REPLACED AND OPERATED NORMALLY. THE OLD RHEOSTAT HAD BEEN REPLACED FEB 2/09.

DURING REPLACE OF THE RT ENGINE, AFT LOWER COWLING FWD BULKHEAD P/N 101-910038-3, CRACKS WERE DISCOVERED IN THE LT AND RT COWLING AFT DUCTS P/N 101-910049-11 AND 101-910049-12 UPPER MOUNTING FLANGES. THIS AREA IS ONLY ACCESSIBLE BY REMOVING THE PANEL (DOOR) PN 101-910049-47 LOCATED IN THE CTR OF THE AFT COWL DUCT ASSY AND USING A BRIGHT LIGHT AND A MIRROR OR BY USE OF A BORESCOPE. THIS PROBLEM WAS DISCOVERED ON A HIGH TIME ACFT WITH 9010 HOURS TT IN SERVICE, BUT THE CONDITION COULD BE PRESENT ON LOWER TIME ACFT. RECOMMEND THIS AREA BE INSPECTED DURING A 'DETAILED ENGINE INSP' SUCH AS A PHASE 4 OF THE ACFT RECOMMENDED INSP PROGRAM.

DURING TROUBLESHOOTING OF INOPERATIVE CABIN COLD CATHODE ("FLUORESCENT") LIGHTING SYS, FOUND SEVERAL INCORRECT STANDARD BUTT SPLICE TYPE CONNECTORS IN HIGH VOLTAGE WIRING SPLICES. MFG COMMUNIQUÉ NR 28 AND OTHERS REFER TO THIS SUBJECT. HAVE PREVIOUSLY FOUND INCORRECT SPLICES DESIGNED FOR LOW VOLTAGE DC WHEN USED IN SHIELDED HIGH VOLTAGE AC WIRING TO MELT OR OVERHEAT AND CAUSE MINOR FIRES ABOVE THE HEADLINER AREA AND/OR CABIN SMOKE. RECOMMEND MFG ADD SPECIFIC WIRE SPLICE REPAIR INFORMATION TO THE APPLICABLE SECTION OF THE MM AND POSSIBLY ISSUE A ONE-TIME MANDATORY SB TO VERIFY CORRECT WIRING SPLICES INSTALLED IN CABIN LIGHTING HIGH VOLTAGE AC WIRING. INCORRECT REPAIRS SEEM TO OCCUR MORE FREQUENTLY WHEN INTERIOR REFURBISHMENT OR REPAIRS ARE ACCOMPLISHED BY FACILITIES OR PERSONNEL UNFAMILIAR WITH SYS REPAIR WIRING USING COMMON SHOP MATERIALS.

(CAN) DURING A POST MAINTENANCE TEST FLIGHT THE LANDING GEAR WAS SELECTED DOWN, THE NOSE AND LEFT DOWN LIGHTS ILLUMINATED AND THE RIGHT STAYED OUT. TWO MORE GEAR CYCLES WERE PERFORMED, HOWEVER THE RIGHT DOWN AND LOCKED LIGHT FAILED TO COME ON. THE PILOT THEN PERFORMED A FLYPAST THE TOWER TO CONFIRM THE GEAR APPEARED TO BE DOWN PRIOR TO CONDUCTING AN UNEVENTFUL LANDING. A POST FLIGHT GEAR SWING WAS PERFORMED AND A LOUD SQUEALING NOISE WAS NOTED COMING FROM THE RIGHT MAIN GEAR ACTUATOR. THE ACTUATOR WAS EXCHANGED FOR ANOTHER UNIT AND THE AIRCRAFT RETURNED TO SERVICE. INSPECTION OF THE ACTUATOR REVEALED THE LOWER HOUSING BEARING WAS IN CONTACT WITH THE PINION GEAR CAUSING IT TO BIND. AN ADDITIONAL NOTE OF INTEREST IS THE ACTUATOR HOUSING WAS FILLED WITH WHAT APPEARS TO BE AEROSHELL 22 INSTEAD OF THE MOLYBDENUM DISULFIDE (MIL-G 21164) THAT IS CALLED UP IN THE COMPONENT MAINTENANCE MANUAL.

PILOT SELECTED FLAPS DOWN THEN NOTICED A ASYMMETRICAL FLAP CONDITION, FLAPS WERE SELECTED UP WHICH RESULTED IN THE CIRCUIT BREAKER POPPING. AIRCRAFT LANDED WITHOUT INCIDENT. MAINTENANCE FOUND THAT THE LEFT OUTBOARD FLAP, INBOARD TRACK HAD FRACTURED NEAR THE TIP CAUSING THE FLAP ROLLER TO COME OUT OF THE TRACK (SEE ATTACHMENT). TRACK REPLACED AIRCRAFT INSPECTED AND RETURNED TO SERVICE.

THE AIRCRAFT DEPARTED FROM YAT AND AFTER REACHING CRUISE ALTITUDE, THE CREW NOTICED SMOKE IN THE CABIN. THE CREW ATTEMPTED TO RETURN TO YAT BUT THE FLAPS WOULD NOT EXTEND ON APPROACH. THE CREW DIVERTED TO YMO. THE FLAP CONTROL CIRCUIT BREAKER TRIPPED, THE SMOKE IN THE CABIN DISSIPATED AND THE AIRCRAFT LANDED SAFELY IN YMO WITH THE FLAPS UP. MAINTENANCE FOUND THAT THE RIGHT FLAPS HAD TRAVELED SLIGHTLY HIGHER THAN NORMAL. WHEN THE FLAP MOTOR AND GEARBOX WAS REMOVED SOME OF THE FLAP ACTUATOR DRIVE CABLES WERE FOUND DAMAGED. IT IS SUSPECTED THAT DUE TO THE EXTREME COLD TEMPERATURES OF THAT DAY, THE UPLIMIT MICROSWITCH DID NOT ACTUATE TO STOP THE FLAPS IN THE UP POSITION. THE FLAP MOTOR THEN CONTINUED TO OPERATE UNTIL THE POINT IT OVERHEATED AND CAUSED THE SMOKE IN THE CABIN AND DAMAGE TO THE DRIVE CABLES. ALL AFFECTED COMPONENTS ARE BEING INSPECTED AND REPLACED AS NECESSARY PRIOR TO RETURNING THE AIRCRAFT TO SERVICE.

DURING THE COMPLIANCE WITH RECOMMENDED SB 27-3929, BOTH ELEVATOR TORQUE TUBES MOUNT BOLTS (8 TOTAL, 4 PER SIDE) WERE SHANKED OUT AND NOT TIGHT AT THE ATTACH POINTS. THE IPC CALLS FOR BOLTS AN-23-10, AND THE REPLACEMENT BOLTS ARE AN23-9. THIS SHOULD BE A MANDATORY SB SINCE IT INVOLVES THE ATTACHMENT OF PRIMARY FLIGHT CONTROLS.

ELT ANTENNA BREAKS IN HALF AND LEAVES THE AIRPLANE IN FLIGHT. WE HAVE INSTALLED (2) ELT ANTENNAS P/N 110-323 AND BOTH HAVE BROKEN OFF. WE ARE NOW INSTALLING A THIRD. THEY KNOW THERE IS A PROBLEM WITH THE ANTENNAS, BUT AT THIS TIME, IT IS NOT A PRIORITY TO CORRECT THE ISSUE. ANOTHER ANTENNA TO REPLACE THE CURRENT ANTENNA, THE PAPERWORK TO ALLOW FOR THE NEW ANTENNA TO BE INSTALLED IS NOT A PRIORITY AT THIS TIME.

THE POWER CABLE WAS REPLACED 200 HRS PRIOR DUE TO VERY STIFF OPERATION AND 200 HRS LATER DURING A SCHEDULED INSPECTION. THE SAME CABLE WAS FOUND VERY STIFF AND HAD EXCESSIVE END PLAY. SUSPECT SUB-QUALITY AIRCRAFT PART.

(CAN) WHILE ACFT WAS IN FOR MAJOR OVERHAUL AND PAINT, THE ELEVATOR BELLCRANK TORQUE TUBE WAS REMOVED FOR ROUTINE MX. THE BELLCRANK ARMS WERE FOUND LOOSE. AFTER THE ARMS WERE REMOVED THE STEEL TORQUE TUBE WAS NDT INSPECTED. CRACK INDICATIONS WERE FOUND ON TWO (2) OF THE EIGHT (8) TAPER PIN HOLES. THIS WAS A NEW ASSEMBLY LESS THAN A YEAR AGO. THE PREVIOUS ELEVATOR BELLCRANK TORQUE TUBE ASSY WAS REPLACED BECAUSE OF CRACKS IN THE SAME AREA. IMPLEMENTING AN INHOUSE INSP OF THIS ELEVATOR BELLCRANK TORQUE TUBE ASSY ON OUR ACFT.

PILOT REPORTED AFTER PUTTING THE GEAR SELECTOR IN THE UP POSITION, THE GEAR WOULD NOT RETRACT. ON TROUBLESHOOTING, THE MECHANIC CYCLED GEAR SEVERAL TIMES AND GOT GEAR RETRACTION TO FAIL. REPLACED RELAY WITH NEW IAW DATE CODE MFD RW 0734. NO RECOMMENDATIONS AT THIS TIME. (K)

WHILE WORKING ON THE AIRCRAFT IN THE HANGAR, IT WAS DISCOVERED THAT THE RT SLIDING DOOR WOULD ONLY OPEN APPROX SIX INCHES. IT WAS FOUND THAT ONE OF THE UPPER "D" RINGS ON THE RIGHT REAR BULKHEAD HAD ROTATED (PROBABLY WHEN THE CARGO NET WAS REMOVED FOR MAINTENANCE) AND IT WAS PREVENTING THE DOOR FROM OPENING FURTHER. THE RING WAS ROTATED TO ITS PROPER ORIENTATION AND THE DOOR NOW OPERATES NORMALLY.

INTERMEDIATE GEARBOX RECEIVED TO INVESTIGATE HIGH ROTATIONAL FORCE TO TURN INPUT PINION. MEASURED TO BE 50 INCH POUNDS. THIS GEARBOX WAS INSTALLED ON AIRCRAFT IN 1995 AND UNTIL NOVEMBER OF 2008 HAD NEVER BEEN REMOVED. DISASSEMBLED GEARBOX AND FOUND AND ACCUMULATION OF MUD LIKE DEPOSIT INSIDE GEARBOX. THIS DEPOSIT WAS ON THE GEARS AND ALL THE BEARINGS WERE COVERED WITH IT. ALL COMPONENTS WERE THEN CLEANED IN WHITE MINERAL SPIRITS. VISUALLY UNDER INSPECTION WITH A BEARING SCOPE THE BEARINGS LOOKED GOOD. BUT UNDER PRESSURE THE DUPLEX BEARING HALVES WERE VERY DIFFICULT TO TURN. THE BEARINGS WERE TURNED OVER TO BELL HELICOPTER PRODUCT SUPPORT FOR FURTHER INVESTIGATION. AT THIS TIME I HAVE NO RESPONSE AS TO WHY THE BEARINGS WERE SO TIGHT. THE GEARBOX WAS THEN UPGRADED TO OVERHAUL AND BOTH DUPLEX BEARINGS AND ALL SEALS WERE REPLACED. AFTER REBUILD THE ROTATIONAL FORCE TO TURN PINION WAS 4 INCH POUNDS.

SEVERAL CRACKS DISCOVERED WHERE SUPPORT ASSEMBLY, ATTACHES TO THE FORWARD SIDE OF THE CROSS TUBE FUNNEL. ONE CRACK EXTENDS PAST THE STANDARD REPAIR AREA AS STATED IN THE SRM. OTHER CRACKS ARE OUT SIDE THE ALLOWABLE REPAIR AREA ALONG THE 90 DEGREE BEND AT THE BOTTOM OF THE SUPPORT P/N 206-031-301-023. NEW SUPPORTS INSTALLED ON THE AIRCRAFT.

THE AIRCRAFT WAS BEING POSITIONED IN THE HANGER FOR MAINTENANCE. THE TAIL SKID WAS BEING USED TO STEADY THE HELICOPTER AS IT WAS BEING HOISTED TO HAVE THE LANDING GEAR REMOVED. A PORTION OF THE TAIL SKID BROKE OFF IN THE ENGINEER’S HAND. THE CAUSE OF THE BREAK IS DUE TO CORROSION. A SERVICEABLE UNIT WAS INSTALLED.

(CAN) APPROX 100 HRS PRIOR TO SCHEDULED REMOVAL OF THE FREEWHEEL ASSY, METAL WAS DISCOVERED ON FREEWHEEL CHIP DETECTOR. CREW ORDERED REPLACEMENT ASSY AND MONITORED FURTHER INDICATIONS. UPON DISASSEMBLY AT THE OPERATOR`S OVERHAUL FACILITY, IT WAS NOTED THAT THE CLUTCH CAGE HAS BEEN CRACKED THROUGH FWD AND AFT FACES. THE "DOG BONE" POCKETS EXHIBITED UNUSUAL WEAR ON INNER SIDE AS WELL AS FWD AND AFT SURFACES. THIS O/H SHOP HAS NOTED SEVERAL CASES OF WORN POCKETS (WEAR BEYOND .004 INCH) ON CLUTCHES INSPECTED WITHIN THE LAST YEAR.

DURING A FUEL CELL CHANGE (DUE TO FUEL SEEPAGE), A SMALL CRACK WAS FOUND IN THE AFT CROSSTUBE SUPPORT STRUCTURE, ON THE RT SIDE OF THE AIRCRAFT. THE LOCATION OF THE CRACK WAS AT THE FWD BOTTOM RIVET OF THE CROSSTUBE SUPPORT. THE CRACK WAS APPROXIMATELY 1/4 INCH IN LENGTH. THIS AREA IS SUSCEPTIBLE TO CRACKS BECAUSE OF THE LOADING ON THIS AREA. THE MANUFACTURER HAS INCREASED THE WALL THICKNESS OF THE SUPPORT STRUCTURE IN THIS AREA, WITH THEIR REPLACEMENT PARTS. THE SUPPORT STRUCTURE WAS REPLACED WITH THE NEW BEEFIER PARTS, AND RETURNED TO SERVICE. IT WAS NOTED THAT THE CRACK COULD NOT BE SEEN FROM THE OUTSIDE, AND THAT ANOTHER CRACK WAS ALSO FOUND ON ANOTHER PIECE OF THE STRUCTURE, WHICH COULD NOT BE SEEN AS WELL. IT WAS ONLY WHEN THE FUEL CELL WAS REMOVED, THAT YOU COULD SEE THE CRACK STARTING FROM THE RIVET.

(CAN) DURING NORMAL FLAT AND LEVEL CRUISE FLIGHT AT APPROX. 4000` ASL THE PILOT INITIATED A GENTLE RT TURN. UPON MOVING THE CYCLIC CONTROL THE PILOT NOTED ROUGH OPERATION AND FEEDBACK THROUGH THE CYCLIC GRIP. THE PILOT CUT HYD AND RESTORED THEM. THE CYCLIC MOVED SMOOTHLY FOR APPROX. 10 SECONDS AND THEN RESUMED ROUGH OPERATION. THE PILOT ONCE AGAIN CUT AND RESTORED HYD. THE CYCLIC MOVED FREELY FOR APPROX. 5 SECONDS AND THEN RESUMED ROUGH OPERATION. THE PILOT ONCE AGAIN CUT HYD BUT THIS TIME RETURNED TO BASE HYD OFF. UPON INSP, ON THE GROUND IT WAS NOTED THAT THE HYD FILTER`S CLOGGING INDICATOR HAD "POPPED". THE FILTER WAS REMOVED FOR VISUAL INSP AND IT WAS DISCOVERED TO BE COMPLETELY CONTAMINATED WITH WHAT APPEARED TO BE BRASS SHAVINGS. THE SOURCE OF THIS MATERIAL IS KNOWN TO BE THE HYD PUMP. THERE WAS NO CAUTION LIGHT THAT INDICATED FAILURE FOR THIS DEFECT AS THE PUMP WAS STILL FUNCTIONING AS IT FAILED. THE SERVOS WERE ESSENTIALLY STARVED OF HYD FLUID PRESSURE DUE TO THE CLOGGED FILTER. THIS PARTICULAR PUMP ONLY HAD A TSN OF 6.5 HOURS.

CUTOFF VALVE ASSY FAILED. THE EPOXY BOND THAT RETAINS THE END CAP IN THE VALVE FAILED. THEN FUEL PRESSURE WAS PRESENT IN THE VALVE ASSY THE PRESSURE WOULD PUSH THE END CAP PARTIALLY OUT OF THE VALVE ASSY. THIS EFFECTIVELY INCREASED THE LENGTH OF THE VALVE. THIS INCREASED LENGTH BLOCKED THE CUTOFF SEAT FOR AN ADDITIONAL 26 THROTTLE DEGREES, CHANGING THE FUEL FLOW CUTOFF ANGLE FROM 8-14 THROTTLE DEGREES TO APPROX 40 THROTTLE DEGREES. THIS ALSO ELIMINATED ALL START FUEL FLOW MODULATION. ENGINE WOULD NOT START. (K)

CORROSION OF LONGERON AND SPLICE UNDER SPLICE. FOUND WHILE PERFORMING INSPECTION REQUIRED FOR ASB212-90-63. CORROSION WAS VERY SEVERE WHEN SIGNS WERE VISIBLE BECAUSE CORROSION WAS BETWEEN LONGERON AND SPAR SPLICE FAYING SURFACES.

(CAN) SUPPORT ANGLE P/N 407-023-800-127 FOUND CRACKED ON A DAILY INSP. PART HAS BEEN UPGRADED BY A -129 THAT WAS INSTALLED. THIS IS KNOWN TO MFG AND THEY HAVE COME OUT WITH THE -129 THAT IS THINNER MATERIAL AND HAS MORE FLEXABILITY.

DURING INSPECTION CABLE TOP END OF MANUAL RELEASE CABLE WAS FOUND TO BE FRAYED AND PULLED OUT OF THE SWAGED END. SDR WAS SUMITTED DEC 09 2008 FOR A SIMILAR DEFECT BUT ON THE BOTTOM END OF CABLE, THAT CASE RESULTED IN A UN-COMMANDED RELEASE OF THE HOOK, BOTH CABLES HAVE BEEN SENT TO THE MANUFACTURE.

DURING A DAILY INSPECTION, THE AFT FAN SHAFT HANGAR BEARING SUPPORT BRACKET WAS FOUND TO BE BROKEN. THERE WAS NO DAMAGE TO THE SURROUNDING AREA, AND INSPECTION OF THE FORWARD BRACKET FOUND NO DAMAGE TO IT. A ONE TIME INSPECTION OF THE WHOLE AIRCRAFT WAS ALSO PERFORMED.

DURING ANNUAL INSPECTION, FRETTING RESIDUE WAS NOTED ON BOTTOM OF LEFT HAND PYLON BEAM ASSEMBLY 427-010-201-107. FURTHER INVESTIGATION AND SUBSEQUENT REMOVAL OF THE BEAM REVEALED CHAFING DAMAGE TO PYLON BEAM 427-010-201-107, ADAPTER 427-010-206-101 AND HOSE 427-365-281-101. THE CAUSE OF THE DAMAGE WAS SLEEVE 120-067B.

ON GROUND RUN IT WAS DISCOVERED THAT CYLINDER NR 5 ON THE RIGHT ENGINE, EGT WAS NOTICABLY HIGHER THAN OTHER CYLINDERS. INDUCTION TUBE WAS REMOVED FROM THE CYLINDER, THE INDUCTION TUBE SLEEVE IN THE OIL SUMP WAS FOUND TO BE LOOSE.

ON GROUND RUN, RT PROP WOULD NOT FEATHER. OTHER PITCH CHANGE FUNCTIONS WORKED NORMALLY. INITIALLY THE PROP WAS REPLACED WITH SIMILAR RESULTS THE GOVERNOR WAS THEN REPLACED WITH A SERVICEABLE UNIT FROM ANOTHER AIRCRAFT AND THE PROP PITCH/FEATHER FUNCTIONED NORMALLY. THE SUSPECT GOVERNOR WILL BE SENT FOR ANALYSIS AND LIKELY OVERHAUL.

IFSD, ACFT TURNED BACK INTO ATLANTA DUE TO, HIGH VIB, COMPRESSOR STALL, AND TGT WENT TO 1149 DEG. FOUND METAL NUGGETS IN THE TAIL PIPE, AND HPT CASE BULGED IN LINE WITH HPT1. ENGINE WILL BE INDUCTED TO MANUFACTURER ON MARCH 3 FOR INVESTIGATION. FURTHER INFORMATION WILL BE AVAILABLE FOLLOWING ENGINE DISASSEMBLY.

(CAN) SURGE/IFSD/AIR RETURN. DURING CLIMB OUT OF DEPARTURE AIRPORT AT 5000 FEET, LOUD BANG REPORTED ACCOMPANIED BY ACFT YAW AND ENG L COMPRESSOR STALL ALERT. ENGINE AUTO SHUTDOWN. ACFT RETURNED TO DEPARTURE. INITIAL FINDINGS NOT AVAILABLE AS OF THIS WRITING. ENGINE HAD BEEN SCHEDULED FOR REMOVAL TONIGHT (10 FEB) FOR HIGH-TIME LIP5 HP TURBINE BLADES. ENGINE WILL BE INDUCTED AT MFG, FURTHER DETAILS WILL BE SUBMITTED AFTER ENGINE TEAR DOWN.

APPROXIMATELY 3.5 HOURS INTO THE FLIGHT TEST PROFILE, CLIMBING OUT OF FL380 TO FL410 AT CLIMB THRUST, IN SMOOTH AIR CONDITIONS, A SUDDEN YAW AND VIBRATION WAS FELT IN THE AIRPLANE. ENGINE INSTRUMENTATION OF THE ENGINE NR 1 SHOWED EGT WAS IN THE YELLOW AT ABOUT 500 C, N2 WAS 70 PERCENT AND N1 WAS AT 30 PERCENT. FLIGHT CREWMEMBERS DECIDED TO SHUTDOWN THE ENGINE, NOTIFIED ATC AND RETURN TO THE BASE, NO EMERGENCY WAS DECLARED. AFTER LANDING, AN INITIAL INSPECTION OF THE NR 1 ENGINE EXHAUSTS AREA SHOWED MISSING TURBINE TIP BLADES. THE ENGINE IS CURRENTLY BEING REMOVED AND WILL BE SENT FOR INVESTIGATION/REPAIR.

DEPARTING YFB FOR YOW ON JAN 31, 2009, THE CREW OBSERVED A NOSE LANDING GEAR FAIL TO LOCK CONDITION AFTER GEAR WAS SELECTED UP. THE AIRCRAFT RETURNED TO POINT OF DEPARTURE AND LANDED WITHOUT FURTHER PROBLEM. MAINTENANCE SERVICED THE NOSE GEAR AND THE AIRCRAFT WAS RETURNED TO SERVICE AND THE AIRCRAFT RETURNED TO YOW.

DURING ROUTINE SCHEDULED HEAVY MAINTANANCE VISIT (C3 CHECK), CORROSION WAS FOUND ON A LAP JOINT AT STA 460 TO 480, ON STR 26L. THIS CORROSION IS BELIEVED TO BE CAUSED BY THE BONDING PROCESS AT FACTORY. THE PROCESS (COLD BONDING) WAS ACCOMPLISHED ON THE FIRST 491 B737 AIRCRAFT. THIS AIRCRAFT IS LINE NR 4481. THIS IS A KNOWN FLAW, AND HENCE THE CHANGE TO A NEW PROCESS ON AIRCRAFT LINE NR 492 AND ON. THIS CORROSION WAS CLASSIFIED, AND VERIFIED AS LEVEL 2, IAW THE CPCP DOCUMENT D6-38528, FIGURE 9 LOGIC DIAGRAM, ITEM NR 20. DUE TO THE LINE NR OF THIS AIRCRAFT THE CPCP PROGRAM HAS SOME TASKS DUE AT 18 MONTH INTERVALS INSTEAD OF MOST AIRCRAFT AT 2 YEAR INTERVALS. AN ADDITIONAL INTERNAL DVI INSPECTION WILL BE RAISED AGAINST THIS AIRCRAFT FOR THIS AREA AND ADJACENT AREAS ONLY AT NEXT HEAVY MAINTENANCE VISIT. THIS WILL BE IN ADDITION TO THE 18 MONTH CPCP INSPECTION AND THE ADDITIONAL INSPECTIONS THAT WILL BE REQUIRED DUE TO THE HMV VISIT.

ON FEBRUARY 8, 2009 LEAVING TORONTO THE AIRCRAFT FAILED TO PRESSURIZE AFTER TAKEOFF THEN RETURNED TO GATE. MAINTENANCE ACTION CONFIRMED THE L1 ENTRY DOOR GUIDE ARM JAM NUT WAS LOOSE CAUSING THE DOOR TO MIGRATE OUT OF RIG. THE GUIDE ARM WAS RE-ADJUSTED. WEST JET CONSIDERS THIS SDR CLOSED.

ON TAXI OUT, R RUDDER TEST FAILED. RUDDER PEDAL APPEARS TO BIND. UNABLE TO ACHIEVE MORE THAN HALF DEFLECTION. INITIAL INSPECTION FOUND CTR PCA (LT HYD SYS) FULLY RETRACTED IN THE LT RUDDER DEFLECTION POSITION, WITH ONLY LT HYD SYS PRESSURIZED, RUDDER APPROX 4 INCHES RIGHT OF NEUTRAL PCA REMAINED FULL LEFT. RIGHT RUDDER PEDAL INPUT APPROX 1/2 TRAVEL, PCA SUDDENLY WENT FULL RIGHT THEN BACK TO FULL LEFT AND REMAINED IN THAT POSITION FOR REMAINDER OF TEST. WITH ALL 3 HYD SYSTEMS PRESSURIZED, FULL RIGHT PEDAL RUDDER MOVED TO THE RIGHT. LEFT RUDDER PEDAL WAS STIFF TO OPERATE AND RUDDER MOVED TO THE LEFT WITH SOUNDS OF BREAKING COMPOSITE. LT HYD SYS PCA JAMMED IN FULL LEFT WITH CTR AND RIGHT SYS PCA`S FIGHTING THE LEFT. ALSO, FOUND INTERFERENCE DAMAGE TO SECONDARY INPUT CONTROL ROD AND REACTION LINK AT CTR PCA POSITION.

INFO RECEIVED FROM MX, WHEN MX HAD BOARDED THE ACFT AFTER ARRIVAL, FRONT END CREW HAD DEPARTED THE ACFT. MX REPORTED THE ACFT AND SPECIFICALLY THE FLIGHT DECK WAS FILLED WITH SMOKE. RELATED SBS: HAMILTON STANDARD 4100941-21-4 OCT 24 2008. THIS ONE IS APPLICABLE TO THE GIVEN P/N BOEING SB 777-21-0104 REV.1 OCT 9 2008. THIS ONE IS APPLICABLE TO P/N 4100943/A/B/C, BUT DESCRIBES THE SAME ISSUE. THE ISSUE IS THE ROTOR RUBBING AGAINST THE STATOR.

(CAN) DURING CRUISE, THE ROTOR BRAKE HANDLE DOWN WARNING LIGHT TURN ON INTERMITTENTLY. THE WARNING LIGHT SWITCH IS ADJUSTED TO WARN PILOT IF THE HANDLE IS NOT IN ITS NEUTRAL OR DOWN POSITION AND TO VERIFY IF THE REASON IS NOT THE PILOT`S SEAT BELT BEING UNDER THE HANDLE. IAW THE MM, THERE SHOULD BE A CLEARANCE OF 22 TO 28 MM BETWEEN THE HANDLE AND THE FLOOR WHEN NOT ENGAGE. A NEW ROTOR BRAKE HANDLE RUBBER BOOT, P/N: 10.2131.9.3, WAS INSTALLED DURING THE LAST O/H, REPLACING THE LEATHER BOOT, P/N: 105-10551, INITIALLY INSTALLED IN ALL MBB BO-105. THE NEW RUBBER BOOT, BEING MORE RIGID THEN THE LEATHER BOOT, KEEPS THE HANDLE UP AT 34MM CLEARANCE BETWEEN THE HANDLE AND THE FLOOR. DURING THEIR PRE-START CHECK, THE PILOT CHECKS IF THE HANDLE IS ALL THE WAY DOWN. BY DOING SO, THEY FORCE THE HANDLE AGAINST THE WARNING LIGHT MECHANISM ON A GREATER DISTANCE AND GET THE SWITCH OUT OF ADJUSTMENT, MAKING THE CONTACTS CLOSING AND TURNING THE ROTOR BRAKE WARNING LIGHT ON WHEN FLYING IN LIGHT TURBULENCES.

LOSS OF PRESSURIZATION AT ALTITUDE ACFT DEPARTED AT FL410, THE CREW HEARD A LOUD NOISE IN THE BACK OF THE ACFT. PRESSURIZATION RAPIDLY LOST. CREW DECLARED EMERGENCY AND DIVERTED TO FLL. UNEVENTFUL LANDING MADE IN FLL. ACFT AT BAS FLL FACILITY. UPDATE, DEADHEAD TRIP FROM NAPLES TO TULSA, OK. WHEN EVENT OCCURRED, CREW DONNED O2 MASKS, MADE SELECTION ON AUDIO CONTROL PANEL AND INITIATED AN EMERGENCY DESCENT. CABIN ALTITUDE DID REACH 15,000 FT. CREW COMMENTED THEY HAD DIFFICULTY WITH COMMUNICATIONS WITH MASKS ON. ALL CABIN MASKED DEPLOYED. NO FAULTS FOUND IN MDC. REQUEST IASCS NVM TO BE DOWNLOADED. A VISUAL INSPECTION OF SAFETY VALVES IN BAGGAGE COMPARTMENT C/W. NO FOREIGN OBJECT DAMAGE FOUND. VISUAL OF OUTFLOW VALVE DONE, NO DEFECTS NOTED.

C/B A-112 FOUND POPPED DURING AIRCRAFT SERVICING. WIRES CONNECTED TO PLUG A3005P1 FOR HOT LIQUID CONTAINER NR 1 WERE FOUND ROUTED SUCH THAT CONTACT WAS MADE BETWEEN TERMINALS CAUSING A SHORT POPPING THE C/B. THE WIRES WERE REPOSITIONED TO PREVENT CONTACT. ADJACENT CONNECTOR A3006P1 FOR HOT LIQUID CONTAINER NR 2 WAS EXAMINED AS WELL AND THE SAME CONDITION EXISTED. THESE WIRES WERE ALSO REPOSITIONED TO PREVENT SHORTING. REFERENCE NR 8 (B) MANUFACTURE IS IN FACT MAPCO. THIS INFO NOT AVAILABLE IN DROP DOWN.

CHIEF PILOT REPORTS DESCENDING BELOW 13000 FT MSL INTO SAN ANTONIO INT`L AIRPORT FROM HOUSTON, THE LEFT ENGINE SHUTTERED, THE HIGH VIB INDICATED ON THE EICAS, THE ITT ROSE TO 700C AND THE OIL PRESSURE DROPPED ABOUT 15 PSI, FROM NORMAL. HE SHUT THE LEFT ENGINE DOWN. THE CREW DECLARED AN EMERGENCY AND LANDED THE AIRCRAFT WITHOUT INCIDENT ON THE RIGHT ENGINE ONLY. ENGINE OIL LEVEL IS REPORTED AS NORMAL, AFTER LANDING. SOME METAL HAS BEEN REPORTED ON THE CHIP DETECTOR OF THE NR 4 AND NR 5 AFT BEARING CASE SUMP RETURN. ALL OTHER CHIP DETECTORS REPORTED TO BE CLEAN OF DEBRIS. ENGINE TO BE REMOVED FOR REPAIR/INVESTIGATION AND LOANER INSTALLED.

WHILE DESCENDING THROUGH 12000 FT ALTITUDE THE CREW TRIED TO START THE APU, BUT THEY EXPERIENCED A HUNG START. THEY THEN SHUT THE ENGINE DOWN AND WAITED TWO MINUTES BEFORE ATTEMPTING A 2ND START WHICH PROVED TO BE SUCCESSFUL. ON LANDING, THEY NOTICED THAT THE APU INTAKE HAD EXPERIENCED SOME KIND OF BACKFIRE WHICH MUST HAVE HAD FLAMES BLOWING OUT THE INTAKE, AND AS A RESULT BURNT THE AREA AROUND THE APU INTAKE. THE INLET DUCT HAS ALSO SUSTAINED SUBSTANTIAL DAMAGED, IN THAT THE COATING IN THE DUCT HAS BURNT AWAY, AND THE FIBERGLASS FIRE PROOF MATERIAL IS NOW EXPOSED. A BOROSCOPE INSPECTION OF THE APU FOUND CHUNKS OF CARBON WEDGED INTO THE TURBINE INLET STATOR. THERE IS ALSO MAJOR BUILDUP OF CARBON IN THE COMBUSTION LINER. REPAIRS WILL BE CARRIED OUT BY A MAINTENANCE FACILITY IN SOUTH AFRICA.

(CAN) THE LANDING GEAR FAILED TO EXTEND WHEN GEAR LEVER WAS SELECTED DOWN. A MISSED APPROACH WAS CARRIED OUT. AT A SAFE ALTITUDE (3000 FT) THE QRH CHECKLIST WAS INITIATED. THE ALTERNATE LANDING GEAR SYS WAS USED. LANDING GEAR EXTENDED (3 GREENS) AND A NORMAL APPROACH AND LANDING WAS EXECUTED. LINE MX INVESTIGATION CONFIRMED FAILURE OF THE LANDING GEAR SELECTOR VALVE. VALVE ASSY REPLACED AND LANDING GEAR FUNCTIONALLY TESTED IAW THE ACFT`S MM. ACFT RELEASED BACK TO SERVICE WITHOUT FURTHER INCIDENT.

DURING A GROUND MAINTENANCE INSPECTION, THE LT MAIN GEAR STABILIZER BRACE WAS FOUND TO BE MISSING APPROXIMATELY 120 DEGREES OF BRACE BEARING SUPPORT AT THE OUTBOARD HINGE POSITION. PART VENDOR IS GOODRICH.

(CAN) AIR TURN BACK DUE TO POWERPLANT MESSAGE DURING CLIMB. CREW ELECTED TO RETURN TO DEPARTURE AIRPORT WITH NO FURTHER ISSUES. ON GROUND TROUBLESHOOTING DISCOVERED PEC FAULTS. PEC REPLACED AND ACFT FUNCTION TESTED ON RETURNED TO SERVICE.

(CAN) SHORTLY AFTER DEPARTURE, THE CREW OBSERVED A PORT HYD PUMP INDICATION LIGHT WHICH REMAINED ILLUMINATED WITH THE MAIN SYS CYCLING EVERY 5 SECONDS. THE ACFT RETURNED TO POINT OF DEPARTURE AND LANDED WITHOUT FURTHER PROBLEM. MX IDENTIFIED THAT THE STARBOARD FLOW INDICATION SWITCH WAS LEAKING. THE UNIT WAS REPLACED AND THE ACFT RETURNED TO SERVICE.

DOWEL SHEARED IN FLIGHT, THE ACFT WAS A TOTAL LOSS, NO INJURIES TO CREW, BUT WAS CLASSIFIED AS AN INCIDENT. NO PROP STRIKES EVER RECORDED, INCLUDING THE NEW DEFINITION. THE INSURANCE COMPANY REQUESTED THE DOWEL FOR METALLURGY TESTS. GIVEN THE TIMES BETWEEN THE O/H. CALENDAR - ALMOST 11 YEARS AND FLIGHT 1881, DIFFICULT TO PROVIDE ANY PREVENTATIVE ACTION FOR THIS TYPE OF OCCURRENCE. HAVE OWNED ACFT SINCE 1988. ENGINE HAS BEEN O/H TWICE IN THIS TIME. (K)

DURING A ROUTINE 200-HR INSPECTION, CRACKS WERE DETECTED ON THE REAR FIN ATTACH BRACKET THAT IS RIVETED TO THE AFT SPAR OF THE HORIZONTAL STAB. ONE HALF INCH CRACK WAS ALONG THE WELD AT THE TOP FORWARD BOXED AREA AND THE OTHER CRACK WAS PROPAGATING FROM ONE OF THE RIVET HOLES FOR THE FIN ATTACH NUT PLATES. IT SHOULD BE NOTED THAT THE NUT PLATES WERE PREVIOUSLY REMOVED TO COMPLY WITH AD 80-11-04 AND THAT NO RIVETS WERE INSTALLED IN THE RIVET HOLE IN QUESTION. FIN ATTACH BOLTS WERE SECURED WITH NUTS AS PER THE AD. WE OPERATE A LARGE FLEET OF CESSNA 152S AND HAVE REPLACED THIS BRACKET ASSEMBLY ONLY TO FIND IT CRACKING AGAIN IN THE SAME AREA. CESSNA NEEDS TO REDESIGN THE BRACKET AND BEEF IT UP.

UPON PULLING THE NOSE WHEEL FAIRING. NOTED THE NOSE FORK HAD A DISTINCT, REARWARD BEND TO IT. THE CENTERLINE OF THE AXLE, RELATIVE TO THE CENTERLINE OF THE CHROME STRUT TUBE WAS APPROX 1.5 INCH AFT. THE NOSE FORK WAS REPLACED IAW MFG SK182-34C, AND ALL THE NOSE GEAR COMPONENTS AND FIREWALL WERE CAREFULLY INSPECTED FOR DAMAGE, BUT AMAZINGLY, NONE WAS FOUND. THE EXCEPTION WAS THE NOSEWHEEL FAIRING, WHICH WAS IN POOR CONDITION, LITERALLY "POP RIVETED" TOGETHER. THE NOSE WHEEL WAS FITTED WITH WHAT APPEARED TO BE A NEW TIRE, BUT NOTHING WAS IN THE RECORDS. THE OWNER OF THE ACFT WAS NOT AWARE OF ANY INCIDENTS, ALTHOUGH HE HAD NOT OWNED THE ACFT FOR VERY LONG. (ABOUT A YEAR) AND HAD NOT FLOWN IT OVER 5 HRS OF SO. THE ACFT IS BASED AT A SMALL, COUNTRY AIRPORT, AND THE OWNER SUSPECTS SOMEONE "BORROWED" IT , DAMAGED IT, AND REPLACED THE NOSE TIRE TO TRY AND HIDE THE DAMAGE. IT WAS A LOCAL AIRPLANE BEFORE THIS OWNER BOUGHT IT, WITH SAME MECH "MAINTAINING" IT FOR SEVERAL YEARS. (K)

DURING SCHEDULED 500-HR INTERNAL MAGNETO INSPECTION IT WAS DISCOVERED THAT THE ROTOR DISTRIBUTOR PLATE HAD OVERHEATED AND BECOME LOOSE. THE PLASTIC ROTOR BODY WAS MELTED IN THE VICINITY OF THE PLATE DUE TO ARCING CAUSED BY THE LOOSE PLATE. THERE WAS NO OPERATIONAL INDICATION, THE ENGINE RAN NORMALLY WITH NORMAL MAG DROPS.

WHILE PERFORMING SB3-08 ON BOTH LEFT AND RIGHT MAGNETOS LT P/N 4370 LT S/N 08010485 LT TSN 64.7 AND RT P/N 4371 RT S/N 07111952 RT TSO 400.6 THERE WAS EVIDENCE OF IMPROPER WEAR OF THE CARBON BRUSH. THE COIL TAB HAD RESIDUE COATING AND UNACCEPTABLE WEAR MARKS. UNDERCUTTING WAS ALSO NOTED ON THE CARBON BRUSH. IT WAS LATER LEARNED THAT BOTH MAGNETOS HAD THE NEW BRUSH REPLACEMENT AND TERMINATING ACTION IMPLEMENTED. THE AFFECTED MAGNETOS WERE SUBSEQUENTLY REPLACED WITH OTHERS AND WERE SENT OUT FOR REPAIR.

LT FLAP OUTBOARD LOWER ROLLER BROKE COMPLETELY UP CAUSING THE FLAP TO SEIZE INTO THE FLAP TRACK DURING THE FLAPS BEING RETRACTED AFTER TAKEOFF. WITH THE FLAP SEIZE THE FLAP MOTOR CONTINUED TO PULL THE SLAVE FLAP UNTIL THE RETRACT CABLE SNAPPED IN FLIGHT. THE FLAP ROLLER HAS BEEN WELL LUBRICATED EVERY 100 HR OF AIRTIME.

RT FLAP RETRACT CABLE P/N 0510105-94 FAILURE DURING FLAP RETRACTION JUST AFTER TAKEOFF. PILOT DISCOVERED DEFECT DURING FLIGHT VISUALLY AND DID NOT HEAR A FAILURE OF THE CABLE OCCUR IN THE CABIN. CABLE FAILURE CAUSED AIRCRAFT TO ROLL/TURN RIGHT IN FLIGHT DUE TO LT FLAP (SLAVE FLAP) FALLING INTO FULL EXTENDED POSITION AND RT FLAP (MASTER FLAP) FULLY RETRACTING. CONTROL OF THE AIRCRAFT WAS REGAINED ONCE PILOT DISCOVERED THE CAUSE OF THE FLIGHT CONTROL DEFECT. THE U/S CABLE BROKE IN THE AREA WHERE THE CABLE ROLLS OVER THE CENTRAL FLAP PULLEY CLUSTER IN THE MIDDLE OF THE CABIN BEHIND THE HEAD LINER. FURTHER INVESTIGATION IS REQUIRED TO PIN POINT AND EXACT CAUSE OF WHY THIS FAILURE OCCURRED WHICH IS IN PROGRESS. JUDGING FROM MY OWN OPINION IT LOOKS AS IF ONE OF THE 7 OF 19 INTERNAL CABLE BUNDLES HAD FAILED PREVIOUSLY AND THE REST OF THE STANDS LOOK TO HAVE FAILED IN TENSION. THE FAILURE OF THE CABLE DOES NOT OCCUR IN ONLY ONE AREA BUT IN MULTIPLE SPOTS WITHIN APPROX 10 INCH LENGTH. THE AIRCRAFT WILL UNDER GO A FULL ABNORMAL OCCURRENCE INSPECTION AND THE FLAP CONTROL SYSTEM WILL BE DISMANTLED TO INSPECT THE ENTIRE SYSTEM FOR DEFECTS. I WILL KEEP ANY UPDATES POSTED. UPDATE- ONE OF THE LEFT (SLAVE) FLAP ROLLERS FAILED CAUSING THE FLAP TO COMPLETELY SEIZE IN THE FLAP TRACK CAUSING THE MOTOR TO TENSION THE RETRACT CABLE TO THE POINT OF FAILURE. ROLLER P/N 0523921. THE ROLLERS HAVE BEEN LUBRICATED ON THIS ACFT AT THE LEAST EVERY 100 HRS OF AIRTIME AND WE HAVE HAD PAST ISSUES WITH THESE ROLLER FAILING.

AFTER LANDING GEAR WAS RETRACTED IN, IT WAS NOTED THAT THE LT LANDING GEAR ASSY DID NOT FULLY RETRACT INTO THE WELL. THE PILOT WAS INFORMED AND IMMEDIATELY RETURNED TO THE AIRPORT. DURING THE RETRACT TEST, IT WAS FOUND THAT THE LEFT GEAR HUNG UP AT THE MIDWAY POSITION BUT WAS ABLE TO EXTEND BACK DOWN FULLY. UPON REMOVAL OF THE MLG ACTUATOR IT WAS DISCOVERED THAT THE ACTUATOR HSG HAD CRACKED A THE FWD MOST SECURING BOLT HOLE. AFTER FURTHER INVESTIGATION, IT HAS BEEN DETERMINED THAT THIS OCCURRENCE HAS HAPPENED FREQUENTLY TO MANY ACFT IN SAME LOCATION BUT RESULTING IN A GEAR UP LANDING. THE ACTUATOR ASSY IS THE ORIGINAL ACTUATOR THAT CAME WITH THE ACFT AT MFG AND IS DATED 1980. (K)

ONE OF THE NOSE GEAR DOWNLOCK PINS BROKE, AT THE GROOVE MACHINED INTO THE PIN FOR ROLLPIN RETENTION, ALLOWING THE UNSECURED PORTION OF THE PIN TO JAM THE NOSE GEAR DOWNLOCK HOOKS. THIS RESULTED IN A GEAR UNSAFE INDICATION AND SUBSEQUENT NOSE GEAR COLLAPSE ON LANDING.

DURING SCHEDULED 500 HOUR MAGNETO INSPECTION IT WAS NOTED THAT THERE WAS ALOT OF WHITE DUST INSIDE MAGNETO. THEN IT WAS NOTED THAT THE PLASTIC GEAR TEETH WERE PARTIALLY WORN AWAY DUE TO BEING IMPROPERLY MESHED WITH THE PLASTIC DISTRIBUTOR GEAR DUE TO THE GEAR NOT BEING DRIVEN DOWN AND SEATED ON THE MAGNETO SHAFT, BOTH MAGNETOS HAD SAME CONDITION ALTHOUGH ONE WAS FAR WORSE THAN THE OTHER.

DURING 100 HR INSP, FOUND LT AND RT AILERON CABLES WORN THROUGH SEVERAL STRANDS AT WING STA 71. THERE IS A CABLE RUB BLOCK INSTALLED ON THE RIB AT THIS STATION. IT IS CAUSING EXCESSIVE WEAR TO THE CABLES. REPLACED BOTH CABLES IN LT AND RT WINGS. P/N 0510105-362, 0510105-364 AND 0510105-365. RECOMMEND INSPECTING THIS AREA CLOSER SINCE IT IS HARD TO SEE AND THERE ARE NO ACCESS PANELS INSTALLED IN THE AREA. INSPECT THE CABLES LOOKING THROUGH THE FLAP PUSHROD HOLE JUST AFT OF THE REAR SPAR. MOVE THE AILERONS TO FULL UP AND FULL DOWN TO SEE THE AREA OF WEAR ON THE CABLE. WE FOUND THIS ON ANOTHER MODEL AS WELL.

DURING POST INSPECTION, ENGINE FAILED TO START. INTERMITTENT ENGAGEMENT OF BENDIX. ONCE REMOVED AXIAL MOVEMENT OF INTERNAL PARTS AS COMPARED TO A NEW STARTER SEEMED EXCESSIVE. STARTER WAS REPLACED AND NOT FURTHER ISSUES WERE NOTED.

ACFT WAS DOING A STEADY CLIMB AFTER TAKEOFF WHEN HE HEARD A LOUD BANG AND LOST RPM. PILOT LANDED THE PLANE BACK ON THE RUNWAY. FUND THE CYLINDER HEAD ON NR 4 CYLINDER COMPLETELY BLOWN OFF. PROBABLE CAUSE COULD POSSIBLY BE MORE DEFECTIVE CYL ASSY THAN CALLED OUT IN AD 08-19-05. THIS CYL ASSY IS OUTSIDE THE SN RANGE. RECOMMEND INSTALLING CYL ASSY. (K)

AME NOTICED FUEL IN THE DRAIN CAN ON A DAILY CHECK. DURING A SUBSEQUENT GROUND RUN HE FEELS THAT THE ENGINE PERFORMANCE FROM LOW TO HIGH IDLE IS NOT AS FAST AS IT SHOULD BE. THE AIRCRAFT IS GROUNDED AND A REPLACEMENT PUMP ASSY AND SUPPORT PARTS HAVE BEEN DISPATCHED.

THIS BOLT WAS INSPECTED DURING THE ANNUAL INSPECTION OF THIS AIRCRAFT. IT WAS VISUALLY SEEN TO HAVE WHAT LOOKED LIKE A CRACK ON THE HEAD SURFACE OF THE BOLT. THIS COULD BE MISTAKEN AS DAMAGE FROM A DRIFT OR HAMMER DONE DURING INSTALLATION. THE PART WAS REMOVED FROM SERVICE, AND SENT FOR NDT INSPECTION. THE INSPECTOR PERFORMED BOTH MAG PARTICLE AND PENETRANT INSPECTIONS AND DETERMINE WITH BOTH METHODS THAT THE BOLT IS IN FACT CRACKED. THEY ARE GUESSING THE CRACK DEPTH TO BE APPROXIMATELY .020-.030 INCH DEEP. THEY ALSO SUGGESTED THAT THIS COULD HAVE BEEN A MANUFACTURING DEFECT CAUSED DURING ANY HARDENING PROCESS THAT MAY HAVE BEEN PERFORMED. THIS PART IS AVAILABLE FOR FURTHER INSPECTION. THE TIME ON THE PART HAPPENS TO BE THE TOTAL TIME OF THE AIRCRAFT AS THERE IS NO RECORD IN THE LOGS WHICH DATE BACK TO 2001 (THE ONLY LOGS AVAILABLE AT TIME OF REPORT) WHICH INDICATES BOLT HAS BEEN REPLACED.

LANDING GEAR WOULD NOT EXTEND ELECTRICALLY OR BY UTILIZING THE EMERGENCY GEAR EXTENSION SYSTEM. THE ACFT LANDED GEAR UP. POST INCIDENT INVESTIGATION WAS CONDUCTED ON SCENE AND A COMPLETE INSPECTION OF THE LANDING GEAR SYS WAS CONDUCTED WITH NO FINDINGS AS TO THE CAUSE. THE ONLY COMPONENT THAT WAS NOT DISASSEMBLED WAS THE LANDING GEARBOX. SUSPECT THAT FOREIGN MATTER BECAME LODGED IN THE GEARBOX PREVENTING NORMAL OPERATION. HOWEVER, THE GEAR SYSTEM OPERATED NORMALLY BOTH ELECTRICALLY AND MANUALLY AFTER THE INCIDENT. SUSPECT FOREIGN MATTER BECAME DISLODGED DURING INCIDENT FORCES OR SUBSEQUENT HANDLING OF THE AIRCRAFT.

PREVIOUSLY REPORTED TO MFG, DURING THE SUPPLEMENTAL INSP FOR CONTINUED AIRWORTHINESS OF THE AIRPLANE. DURING SUPPLEMENTAL INSP NR 57-10-14 CARRY-THRU FWD SPAR CAP, FOUND MULTIPLE PNEUMATIC LINES SPLICED TOGETHER WITH STANDARD MIL-SPEC HOSE ON LT AND RT WINGS. LINES NOT AVAILABLE FROM MFG. FABRICATED LINES FROM PN R3/4X049-T3, WWT-700/4X.500, WWT-700/4X.625 AND R3/8X035-50. INSTALLED LINES AND UNION BACK TO ORIGINAL IPC 36-10-00 CONFIGURATION.

PERFORMED SUPPLEMENTAL INSP FOR CONTINUED AIRWORTHINESS OF THE AIRPLANE. DURING SUP INSP NR 57-10-14 CARRY-THRU FWD SPAR CAP, FOUND MULTIPLE PNEUMATIC LINES SPLICED TOGETHER WITH STANDARD MIL-SPEC HOSE ON LT AND RT WINGS. LINES NOT AVAILABLE FROM MFG. FABRICATED LINES FROM PN R3/4X049-T3, WWT-700/4X.500, WWT-700/4X.625 AND R3/8X035-50. INSTALLED LINES AND UNION BACK TO ORIGINAL IPC 36-10-00 CONFIGURATION.

PERFORMED SUPPLEMENTAL INSP FOR CONTINUED AIRWORTHINESS OF THE AIRPLANE. DURING SUPPLEMENTAL INSP NR 57-10-14 CARRY-THRU FWD SPAR CAP FOUND MULTIPLE PNEUMATIC LINES SPLICED TOGETHER WITH STANDARD MIL-SPEC HOSE ON LT AND RT WINGS. LINES NOT AVAILABLE FROM MFG. FABRICATED LINES FROM PN R3/4X049-T3, WWT-700/4X.500, WWT-700/4X.625 AND R3/8X035-50. INSTALLED LINES AND UNION BACK TO ORIGINAL IPC 36-10-00 CONFIGURATION.

PREVIOUSLY REPORTED DURING SUPPLEMENTAL INSP FOR CONTINUED AIRWORTHINESS OF THE AIRPLANE. DURING SUP INSP NR 57-10-14 CARRY-THRU FWD SPAR CAP FOUND MULTIPLE PNEUMATIC LINES SPLICED TOGETHER WITH STANDARD MIL-SPEC HOSE ON LT AND RT WINGS. LINES NOT AVAILABLE FROM MFG. FABRICATED LINES FROM PN R3/4X049-T3, WWT-700/4X.500, WWT-700/4X.625 AND R3/8X035-50. INSTALLED LINES AND UNION BACK TO ORIGINAL IPC 36-10-00 CONFIGURATION.

REPORTED TO MFG DURING SUPPLEMENTAL INSPECTION FOR CONTINUED AIRWORTHINESS OF THE AIRPLANE. DURING SUPPLEMENTAL INSPECTION NR 52-10-01 CABIN DOOR RETENTION FOUND HARDWARE UNABLE TO REMOVE FROM UPPER CABIN DOOR DUE TO CORROSION. DRILLED OFF UPPER AND LOWER CABIN DOOR HINGE SUPPORTS TO REMOVE DOOR FOR INSPECTION. INSTALLED TWO NEW HINGE SUPPORTS. REPLACED HARDWARE, BUSHINGS, STAT-O-SEALS, BEARINGS, BOLTS, AND NUTS.

PREVIOUSLY REPORTED DURING SUPPLEMENTAL INSPECTION FOR CONTINUED AIRWORTHINESS OF THE AIRPLANE. DURING SUPPLEMENTAL INSPECTION NR 52-10-01 CABIN DOOR RETENTION FOUND HARDWARE UNABLE TO REMOVE FROM UPPER CABIN DOOR DUE TO CORROSION. DRILLED OFF UPPER AND LOWER CABIN DOOR HINGE SUPPORTS TO REMOVE DOOR FOR INSPECTION. INSTALLED TWO NEW HINGE SUPPORTS PART NUMBER 5111512-8. REPLACED HARDWARE, BUSHINGS, STAT-O-SEALS, BEARINGS, BOLTS, AND NUTS.

PREVIOUSLY REPORTED TO MFG DURING SUPPLEMENTAL INSPECTION FOR CONTINUED AIRWORTHINESS OF THE AIRPLANE. DURING DISASSEMBLY AND INSPECTION IAW SUPPLENTAL INSPECTION NR 32-10-06 FOUND DEEP SCRATCHES AND CORROSION ON THE ISOLATION PISTON PN 5941136-1. REPLACED ISOLATION PISTON.

PREVIOUSLY REPORTED DURING SUPPLEMENTAL INSPECTIONS FOR CONTINUED AIRWORTHINESS FOR THE AIRPLANE. DURING REASSEMBLE OF ELEVATOR TRIM TAB ACUTATOR IAW MM 27-30-02 FOUND BARREL CRACKED AND VERIFIED BY EDDY CURRENT INSPECTION. REPLACED ACTUATOR WITH PN 5715160-8.

REPORTED DURING SUPPLEMENTAL INSPECTIONS FOR CONTINUED AIRWORTHINESS FOR THE AIRPLANE. DURING INSPECTION OF ELEVATOR TRIM TAB ACUTATOR IAW MM 27-30-02 FOUND BARREL CORRODED AND PITTED. REPLACED ACTUATOR WITH PN 5715160-8.

REPORTED DURING SUPPLEMENTAL INSPECTIONS FOR CONTINUED AIRWORTHINESS FOR THE AIRPLANE. DURING INSPECTION OF ELEVATOR TRIM TAB ACUTATOR IAW MM 27-30-02 FOUND BARREL CORRODED. REPLACED ACTUATOR WITH PN 5715160-8.

REPORTED DURING SUPPLEMENTAL INSPECTIONS FOR CONTINUED AIRWORTHINESS FOR THE AIRPLANE. DURING INSPECTION OF ELEVATOR TRIM TAB ACUTATOR IAW MM 27-30-02 FOUND BARREL CORRODED. REPLACED ACTUATOR WITH PN 5715160-8.

PREVIOUSLY REPORTED DURING SUPPLEMENTAL INSPECTION FOR CONTINUED AIRWORTHINESS OF THE AIRPLANE. DURING DIASSEMBLY AND INSPECTION IAW SUPPLENTAL INSPECTION NR 32-10-06 FOUND DEEP SCRATCHES AND CORROSION ON THE ISOLATION PISTON PN 5941136-1. REPLACED ISOLATION PISTON.

DURING THE SID INSPECTION THE RUDDER WAS REMOVED. VISUALLY DISCOVERED THE TOP RUDDER PIVOT BEARING FALLING APART. THERE IS NO REQUIREMENT TO REMOVE THE RUDDER EXCEPT DURING THE SID INSPECTION CALLED OUT IN THE LATEST MANUAL REVISION 16. COULD NOT BE DETECTED WITHOUT RUDDER REMOVAL.

DURING THE SID INSPECTION THE INTERIOR UPHOLSTERY WAS REQUIRED TO BE REMOVED. VISUALLY FOUND THE FUSELAGE DOOR FRAME STRUCTURE WHERE THE AFT DOOR SUPPORT CABLE MOUNTS TO BE CRACKED. THERE IS NO REQUIREMENT TO REMOVE THE SIDE WALL TRIM EXCEPT DURING THE SID INSPECTION CALLED OUT IN THE LATEST MANUAL REVISION 16. THIS COULD NOT BE DETECTED WITHOUT UPHOLSTERY REMOVAL.

DURING THE SID INSP THE RUDDER WAS REMOVED. VISUALLY DISCOVERED THE RUDDER TOP HINGE BRACKET HOLES THAT BOLT TO THE BEARING TO BE EXCESSIVELY ELONGATED. THERE IS NO REQUIREMENT TO REMOVE THE RUDDER EXCEPT DURING THE SID INSP CALLED OUT IN THE LATEST MANUAL REVISION 16. THIS WOULD NOT BE DETECTED WITHOUT RUDDER REMOVAL.

DURING THE SID INSP, THE NLG WAS REMOVED FOR INSPECTION. THE RETRACT ACTUATOR UPPER MOUNT WAS FOUND CRACKED WHICH ALLOWED THE DOWNLOCK OVER-CENTER TENSION TO BE REDUCED. THERE IS NO REQUIREMENT TO REMOVE THE NLG EXCEPT DURING THE SID INSPECTION CALLED OUT IN THE LATEST MANUAL REVISION 16. IT WOULD HAVE BEEN VERY DIFICULT TO FIND THIS PROBLEM WITHOUT REMOVING THE NLG.

DURING THE SID INSP, THE NLG WAS REMOVED FOR INSP. THE DRAG BRACE LT AND RT MOUNT BOLT HOLES WERE ALL FOUND EXCESSIVLY ELONGATED REQUIRING STRUCTURAL REPAIR. SUSPECT DAMAGE WAS CAUSED BY NORMAL TOWING OVER THE 9268 HOURS. THIS WOULD NOT BE DETECTED UNLESS THE NOSE GEAR IS REMOVED AS REQUIRED BY THE SID INSP CALLED OUT IN THE LATEST MANUAL REVISION 16.

DURING THE SID INSP, THE UPPER CABIN DOOR WAS REMOVED FOR INSP. THE HINGE PIVOT BOLTS COULD NOT BE REMOVED BECAUSE THE BOLTS, BUSHINGS AND BEARINGS WERE ALL RUSTED SOLID. THIS CAUSED THE BOLTS TO PIVOT IN THE FUSELAGE STRUCTURE INSTEAD OF THE BEARINGS CAUSING EXCESSIVE WEAR OF THE FUSELAGE DOOR MOUNT HOLES. THERE IS NO REQUIREMENT TO REMOVE THE UPPER CABIN DOOR EXCEPT DURING THE SID INSP, CALLED OUT IN THE LATEST MANUAL REVISION 16.

DURING THE SID INSP, THE LOWER DOOR WAS REMOVED FOR INSPECTION. FOUND LOWER CABIN DOOR AFT CHANNEL SUPPORTING STEP HINGE CRACKING OUT OF ACCESS HOLES. LOWER CABIN DOOR FRAME DOOR HANDLE PIVOT BOLT HOLE IS WORN OUT. CABIN DOOR FRAME IS CRACKING NEAR THE FWD END OF THE LOWER CABIN DOOR END CAP. FOUND THAT THE CABIN DOOR SKIN IS CRACKED.

DURING THE SID INSP, FOUND AREA OF DE-BOND IN HORIZONTAL STABILIZER. REPAIRED DE-BOND AREA BETWEEN PN 5732133-2, STRINGERS AND HORIZONTAL STABILIZER LOWER SKIN PANEL PN 5732130-17 AND -18 IAW FORM 8110-3 INSTRUCTIONS. THIS DE-BONDING WOULD NOT HAVE BEEN FOUND WITHOUT DOING THE SID INSPECTION.

DURING THE SID INSP THE UPPER CABIN DOOR WAS REMOVED FOR INSP. THE HINGE PIVOT BOLTS COULD NOT BE REMOVED BECAUSE THE BOLTS, BUSHINGS AND BRGS WERE ALL RUSTED SOLID. THIS CAUSED THE BOLTS TO PIVOT IN THE FUSELAGE STRUCTURE INSTEAD OF THE BEARINGS CAUSING EXCESSIVE WEAR OF THE FUSELAGE DOOR MOUNT HOLES. THERE IS NO REQUIREMENT TO REMOVE THE UPPER DOOR EXCEPT DURING THE SID INSP CALLED OUT IN THE LATEST MANUAL REVISION NR 16.

DURING THE SID INSP THE WING SURFACE DE-ICE BOOTS WERE REQUIRED TO BE REMOVED. AFTER REMOVAL FOUND LARGE AREAS OF MODERATE CORROSION ON THE RT WING INBD AND OTBD SECTIONS. A SMALLER AREA WAS FOUND ON THE LT WING L/E.

DURING THE SID INSP, THE FUSELAGE BULKHEAD AT THE FWD OPENING OF THE CABIN DOOR AND THE SECONDARY BULKHEAD WERE FOUND CRACKED, BUCKLED AND DISTORTED. THIS WAS ONLY FOUND BECAUSE THE SID INSP REQUIRES REMOVAL OF THE DOOR BAYONET RECEPTACLES FOR INSP AND THE REMOVAL OF INTERIOR SIDEWALL UPHOLSTERY.

DURING THE SID INSP BOTH WINDSHIELDS WERE INSPECTED USING A PRISM, BOTH WINDSHIELDS WERE FOUND CRACKED. THE CRACK CRITERIA IN THE LATEST MANUAL REVISION IS - NO CRACKS ALLOWED. IF USING THE PREVIOUS MANUAL REVISION (OR OLDER REVISION) THESE CRACKS ARE ACCEPTABLE.

WHILE ENROUTE AT 41000 FT HEAVY ACRID BLACK SMOKE THE INUNDATED CABIN. OXYGEN MASKS WERE DEPLOYED AND AN EMERGENCY DESCENT WAS INITIATED. ACFT WAS LANDED AT THE NEAREST SUITABLE AIRPORT WITHOUT INCIDENT. ON FEB 21ST, MX IDENTIFIED THE MALFUCTION AS A OVERHEAD BLOWER FAN AND REPLACED WITH NEW FAN ASSY. THE ACFT WAS SUBSQUENTLY RETURNED TO ITS HOME BASE FOR CLEANING AND INTERIOR RESTORATION.

DURING PREFLIGHT, IT WAS DISCOVERED THAT THE HORIZONTAL STABILIZER TRIM POSITION INDICATION DID NOT CHANGE TO WHITE WHEN TRIM WAS IN TAKEOFF POSITION AREA. NO TAKEOFF MESSAGE REMAINED ON IN THE EICAS AND THE RED STAB NO TAKEOFF ANNUNCIATOR ILLUMINATED WHEN THRUST LEVER TAKEOFF POSITION WAS SELECTED. THE HORIZONTAL STABILIZER TRIM ACTUATOR (HSTA) WAS REPLACED BUT THE REPLACEMENT FAILED THE INDICATION TEST AS THE STAB NO TAKEOFF ANNUNCIATOR WOULD NOT ILLUMINATE. ANOTHER HSTA (OVERHAULED) WAS INSTALLED AND TESTED SERVICEABLE.

WHILE INVESTIGATING A HYD FLUID LEAK IN THE NOSE WHEEL WELL, LOCATED CRACK IN THE RETRACT CYLINDER HEAD BETWEEN THE ATTACH BOLT/BUSHING BOSS AND THE BARREL THREADS, IN THE UPPER RADIUS. THERE WERE NO TOOL MARKS TO CREATE A STRESS RISER, AND NO HARD LANDINGS NOTED. THIS WAS NOT VISIBLE, WITHOUT REMOVING THE RETRACT CYLINDER, AS IT WAS HIDDEN BY THE ATTACH STRUCTURE IN THE WHEEL WELL. APPROX 800 FLIGHT HOURS AND 19 YEARS SINCE LAST CYLINDER REMOVAL/RE-SEAL.

PILOT REPORTED LOW OIL PRESSURE DURING TAKEOFF PHASE OF FLIGHT FOLLOWED BY NO OIL PRESSURE INDICATION SHORTLY AFTER. PILOT THOUGHT THE LOW OIL PRESSURE WAS AN INDICATION PROBLEM. UPON INSPECTION AND FUNCTIONAL CHECK OF PRESSURE GAUGE, THE INDICATOR WAS RULED OUT. A MANUAL GAUGE WAS CONNECTED TO THE ENGINE AND A FUNCTIONAL CHECKED PERFORMED. THE OIL PRESSURE SYS INITIALLY MADE 20 POUNDS PRESSURE FOLLOWED BY A DROP TO AROUND 10 POUNDS WHEN ENGINE RPM WAS SLIGHTLY INCREASED. THE OIL FILTER PN CH48108-1 WAS REMOVED FOR INSPECTION. UPON EXAMINATION THE RELIEF VALVE WAS FOUND DISPLACED AND LOOSE (FLOATING) ABOUT THE FILTER. A NEW FILTER WAS INSTALLED AND A FUNCTIONAL CHECKED PERFORMED AND OIL PRESSURE SYS WAS NORMAL.

DURING SLOW FLIGHT MANEUVERS THE INSTRUCTOR NOTICED THE GREEN GEAR SAFE LIGHT WAS OUT. THE STUDENT OBSERVED THE LEFT GEAR WAS HANGING OUT AND SWINGING. THE GEAR WAS CYCLED BUT THE GEAR REMAINED HANGING. AFTER CONTACTING DISPATCH THE INSTRUCTOR ELECTED TO RETRACT THE RIGHT AND NOSE GEARS AND LAND GEAR UP. MAINTENANCE FOUND THE LEFT GEAR PIVOT HAD SHEARED AT THE SPLINE SHAFT. THE RIGHT GEAR WAS REMOVED AND ZYGLO CHECKED. MAINTENANCE FOUND THE RIGHT PIVOT WAS CRACKED 3/4 AROUND THE SPLINE SHAFT, NEARING COMPLETE FAILURE. THIS CRACK AREA IS NOT NORMALLY ACCESSIBLE DURING REGULAR INSPECTIONS WITHOUT REMOVING THE GEAR ACTUATOR. CESSNA HAS BEEN NOTIFIED OF THE PROBLEM. AT THIS TIME THERE ARE NO REPLACEMENT STEEL PIVOTS AVAILABLE FOR THESE AIRCRAFT.

DURING START AND RUNUP, NR 3 CYLINDER WAS RELUCTANT TO FIRE UNLESS THE MIXTURE WAS LEANED ALMOST TO IDLE CUTOFF. ONCE THE ENGINE WAS WARM IT WOULD RUN NORMALLY. IN FLIGHT THE EGTS WERE UNEVEN, WITH NR 3 AND NR 6 MUCH COOLER THAN THE REST. BOTH OF THESE CYLINDERS FEED OFF THE BOTTOM OF THE PLENUM AND SUSPECT THAT POORLY ATOMIZED FUEL WAS RUNNING DOWN AND ENTERING THEM, MAKING THE MIXTURE DISTRIBUTION VERY UNEVEN. REPLACING THE CARBURETOR ELIMINATED THE PROBLEM ENTIRELY.

DURING ACCOMPLISHMENT OF STC NR ST02664CH, INSTALLATION OF ENGINES, THE PYLONS REQUIRE DISASSEMBLY AND MODIFICATION. DURING DISASSEMBLY OF THE LT PYLON, THE AFT PYLON CENTER RIB WAS FOUND CRACKED AT THE LOWER STRINGER CUT-OUT. THE CRACK EMANATES FWD FROM THE STRINGER CUT-OUT. THE LOWER FLANGE OF THE RIB IS ALSO JOGGLED AT THIS AREA.

DURING ACCOMPLISHMENT OF STC ST02661CH, INSTALLATION OF ENGINES, THE PYLON REQUIRES DISASSEMBLY AND MODIFICATION. DURING DISASSEMBLY OF BOTH LT AND RT PYLONS, THE AFT PYLON CENTER RIB WAS FOUND CRACKED AT THE LOWER STRINGER CUT-OUT. THE CRACKS ARE EMANATING FWD FROM THE STRINGER CUT-OUT. THE LOWER FLANGE OF THE RIB IS ALSO JOGGLED IN THIS AREA.

DURING A 100 HR/ ANNUAL INSP THE AUTOPILOT PITCH TRIM SERVO BRACKET WAS FOUND TO HAVE BROKEN FREE FROM THE 90 DEGREE SUPPORT BRACKET THAT IT ATTACHES TO. THIS BRACKET IS A FACTORY INSTALLED PART FOR THE AUTOPILOT SYS AND IS RIVETED TO THE EXISTING AIRFRAME. THE PROBABLE CAUSE TO THIS FAILURE IS EXCESSIVE VIBRATION IN THIS AREA. RECOMMENDATIONS TO PREVENT A RECURRENCE WOULD BE TO INSTALL A DOUBLER BRACKET ON THE INBD SIDE OF THE ORIGINAL BRACKET TO HELP SHARE THE LOAD AND INCREASE THE SIZE OF THE MATERIAL LOCATED IN THAT AREA. (K)

REMOVED CRACKED LT AND RT WING TIPS AND REPLACED WITH NEW PN`S 0723205-1 AND 0723205-20. REMOVED CRACKED ELEVATOR RT AND LT TIPS AND REPLACED WITH NEW PN`S 1234640-4 AND 1234640-3. REMOVED CRACKED HORIZONTAL STAB PLASTIC TIP ON RT SIDE & REPLACED WITH NEW. ACFT KEPT IN A 50 DEGREE HEATED HANGER. ACFT TAKEN OUTSIDE FOR FLT IN 17 DEGREE TEMP. INFLT TEMPS REACHED -17 DEGREES. UPON LANDING AT DESTINATION, OBSERVED THAT ALL PLASTIC WING, ELEVATOR & HORIZ STABILIZER TIPS SUSTAINED CRACKS. ALL CRACKS WERE OF SEVERITY THAT SEPARATION COULD HAVE RESULTED DURING FLT. ACFT PLACED OUT OF SERVICE UNTIL PARTS COULD BE REPLACED. MFG NOTIFIED. ALL PARTS WERE REPLACED AND THE ACFT WAS PLACED BACK IN SERVICE. (K)

REMOVED CRACKED LT AND RT WING TIPS AND REPLACED WITH NEW PN`S 0723205-1 AND 0723205-20. REMOVED CRACKED ELEVATOR RT AND LT TIPS AND REPLACED WITH NEW PN`S 1234640-4 AND 1234640-3. REMOVED CRACKED HORIZONTAL STAB PLASTIC TIP ON RT SIDE & REPLACED WITH NEW. ACFT KEPT IN A 50 DEGREE HEATED HANGER. ACFT TAKEN OUTSIDE FOR FLT IN 17 DEGREE TEMP. INFLT TEMPS REACHED -17 DEGREES. UPON LANDING AT DESTINATION, OBSERVED THAT ALL PLASTIC WING, ELEVATOR & HORIZ STABILIZER TIPS SUSTAINED CRACKS. ALL CRACKS WERE OF SEVERITY THAT SEPARATION COULD HAVE RESULTED DURING FLT. ACFT PLACED OUT OF SERVICE UNTIL PARTS COULD BE REPLACED. MFG NOTIFIED. ALL PARTS WERE REPLACED AND THE ACFT WAS PLACED BACK IN SERVICE. (K)

NR 4 CYLINDER BARREL CRACKED APPROX. 1INCH FROM THE TOP OF THE BARREL AND ALMOST .7500 INCH OF THE WAY AROUND THE INSIDE CIRCUMFERENCE. THIS WAS A NEW ECI FREEDOM STEEL BARREL CYLINDER. THE CYLINDER HAS TEST CELL AND GROUND RUN TIME ONLY, WHEN THE MECHANIC NOTICED OIL COMING FROM AROUND THE AREA OF THE BOTTOM COOLING FIN OF THE CYLINDER HEAD.

100 HOUR INSPECTION CARRIED OUT AND ENGINE LEAKDOWN CARRIED OUT NORMAL. AFTER ENGINE WAS COOL THE PROPELLER WAS ROTATED TO CHECK MAGNETO TO ENGINE TIMING. A BANG WAS HEARD IN THE FRONT AREA OF THE ENGINE. NR 5 AND NR 6 CYLINDER EXHAUST VALVES WERE STICKING OPEN AND WHEN THE PISTON CONTACTED, THEY WERE JARRED LOOSE AND RETURNED TO THEIR SEATED POSITION WITH VALVE SPRING PRESSURE. ALL CYLINDERS WERE REMOVED AND IT WAS DETERMINED THE EXHAUST VALVE PUSH RODS FOR NR 5 AND NR 1 CYLINDERS WERE BENT, ALONG WITH NR 1, NR 5, AND NR 6 PISTON HEADS HAD SIGNS OF CONTACT WITH THE EXHAUST VALVES. ALL CYLINDERS WERE REPLACED WITH REPAIRED UNITS. NR 1 CYLINDER EXHAUST VALVE WAS REMOVED AND INSPECTED VALVE STEM AND GUIDE. BOTH HAD CARBON DEPOSITS CREATING THE REDUCED CLEARANCE AND STICKING. THE AIRCRAFT HAD BEEN ON SURVEY WORK PRIOR TO INSPECTION WITH REDUCED POWER SETTINGS AND COLD AMBIENT OUTSIDE AIR TEMPERATURES. SUSPECT THIS MAY HAVE CONTRIBUTED SOMEWHAT, BUT THE PILOT REPORTED THAT ENGINE TEMPERATURES WERE NORMAL DURING THIS TIME. THERE WAS NO ROUGH RUNNING ENGINE IN FLIGHT, BUT THE PILOT REPORTED AFTERWARD THAT ONE START-UP ONE MORNING THERE WAS SOME BACKFIRING BUT IT WENT AWAY AND HE ATTRIBUTED IT TO COLD OUTSIDE AIR TEMPERATURES AND POSSIBLE OVER PRIMING.

THE BYPASS DOOR ON THE ENGINE AIR INDUCTION FILTER DUCT TO THE THROTTLE BODY WAS FOUND LAYING IN THE BOTTOM OF THE DUCT, THE DOOR PIN HAD FALLEN OUT AND THE DOOR, THE SPRING, AND THE HINGE PIN WERE FOUND IN THE BOTTOM OF THE DUCT , NO PARTS HAD GONE ANY FURTHER UP THE DUCT. THE BYPASS DUCT DOOR IS LOCATED DOWNSTREAM OF THE FILTER, ALL PARTS WERE ACCOUNTED FOR. THE PIN FELL OUT OF THE HINGE BECAUSE THE HINGE END WAS NOT SUFFICIENTLY STAKED CLOSED.

DURING WINTER MAINTENANCE, A CRACK WAS FOUND AT FUSELAGE STATION 389.0/WING STATION 43 IN THE FUSELAGE SKIN NEAR THE FORWARD LEFT WING/FUSELAGE PICKUP ANGLE ATTACHMENT. THE CRACK IS APPROXIMATELY 2.5 INCH LONG AND RUNS FORE AND AFT ALONG THE BEND RADIUS IN THE FUSELAGE ROOF SKIN.

ACCUMULATOR IS THE ORIGINAL ONE (FLT HRS: 8880.5 AND CYCLES: 6991), THE CREW WAS ONBOARD (NO PASSENGERS) PERFORMING THE PRE FLIGHT AND WHEN SELECTING THE HYDRAULIC THEY HEARD A LOUD BANG. THE EVENT DID HAPPEN AROUND 13:30 AT DETROIT INTERNATIONAL AND THE AIRCRAFT FLEW 3 LEGS BEFORE THE INCIDENT. THE CONNECTOR P3QA INSTALLED ON THE R ENG FUEL SOV (V2QA) WAS HIT AND THE WIRES WERE BROKEN, THEY HAD TO DISCONNECT THE BATTERY SINCE THE WIRES WERE "SPARKING". ONE FUEL FEED LINE DAMAGED, 14TH STAGE BLEED DUCT DAMAGED (PROBABLY BROKEN).

AIRCRAFT TOOK A CREW CANCEL BECAUSE THE FO CALL FATIGUE. THE AIRCRAFT THEN WAS RESCHEDULED TO REPOSITION AT 1000. GOT A CALL THAT THE CAPTAIN’S LEFT RUDDER PEDAL WAS BROKEN. WHEN THE MECHANICS WENT O BOARD THEY FOUND THAT THE CAPTAIN’S RUDDER PEDAL WAS BENT AND TWISTED AND LOOKING DOWN AT THE RUDDER PEDAL TUBE THEY COULD SEE THAT THERE WAS A CRACK.

AIRCRAFT WAS AOG DUE TO A CUT IN MAIN WHEEL TIRE NR 2. SENT A MECHANIC IN ORDER TO REPLACE THE DAMAGED MAIN WHEEL. AFTER REPLACEMENT OF THE MAIN WHEEL, THE MECHANIC WANTED TO PERFORM A TIRE SERVICING CHECK ON ALL TIRES. DURING THE TIRE CHECK THE RT NOSE WHEEL/TIRE EXPLODED. THE RT NOSE WHEEL WAS TOTALLY DESTROYED, BROKEN DOWN IN SEVERAL PARTS. IT HAS BEEN REPORTED THAT THE TIRE ITSELF IS NOT BURST. THE WHEEL BEARING WAS FOUND IN A DISTANCE OF 6 TO 8 METERS FROM THE NLG. A PRESSURE-MEASURING GAUGE WAS ALSO FOUND IN A DISTANCE OF 6 TO 8 METERS FROM THE NLG. IT IS SUSPECTED, THAT THE EXPLOSION OCCURRED DURING/AFTER THE INFLATION AND OR A PRESSURE CHECK OF THE RT NOSE WHEEL. THE MECHANIC WAS HEAVILY INJURED AND LOST ONE ARM AND ONE FOOT/LEG. THE SITUATION OF THE MECHANIC IS CURRENTLY STABLE. HE IS NOT IN DANGER OF LIFE. THE WHEEL ASSEMBLY WAS ONLY SIX DAYS (SINCE DECEMBER 6) INSTALLED ON ACFT AFTER COMING FROM AERO, A LUFTHANSA TECHNIQUE MAINTENANCE FACILITY. THE WHEEL ASSEMBLY HAS THE P/N 5010598 AND S/N SAPR000065 (LT), OCT 930173 (RT). THE TIRES INSTALLED ON THE NOSE WHEEL ARE FROM VENDOR DUNLOP. THE DESTROYED, RIGHT NOSE WHEEL TIRE WAS NEW. THE LEFT NOSE WHEEL TIRE WAS ONE TIME RETREATED. THE DESTROYED, RIGHT WHEEL ASSEMBLY HAS BEEN OVERHAULED AT AUGUST 28, 2008. THE BRITISH AIR ACCIDENT INVESTIGATION BRANCH (AAIB) HAS QUARANTINED THE AIRCRAFT FOR INVESTIGATION AND PRESERVATION OF EVIDENCE.

FLT 5542, IAH-MOB, DURING CLIMB, FLT CREW REPORTED RIGHT ENGINE VIBRATION WAS HIGH AND THERE WAS AN INDICATION OF LOW OIL PRESSURE. RIGHT ENGINE SHUT DOWN AND FLIGHT RETURNED TO IAH AND LANDED WITHOUT FURTHER INCIDENT. MAINT INSPECTED ENGINE AND FOUND A LOOSE B-NUT ON LUBE/SCAVENGE PUMP. B-NUT RESECURED AND OIL WAS SERVICED. ENGINE RUNS CARRIED OUT AND NO DEFECTS NOTED. AIRCRAFT RETURNED TO SERVICE.

FLT CREW REPORTED THAT ON GEAR RETRACTION, THERE WAS A NOSE LANDING GEAR DISAGREE MSG AND THE NOSE LANDING GEAR DOOR INDICATION RED AND THE NOSE LANDING GEAR INDICATES IN TRANSIT. THE AIRCRAFT CONTINUED ON TO IAH AND LANDED WITHOUT FURTHER INCIDENT. MX PERFORMED GEAR SWINGS AND THE GEAR RETRACTED NORMALLY. BUT THERE WERE NO GREEN LIGHTS FOR THE NOSE GEAR DOORS. A PSEU WAS INSTALLED BUT NIL FIX. THE RIGHT NOSE LANDING GEAR PROX SWITCH HAD DAMAGED PINS ON THE HARNESS. PINS WERE REPLACED BY NIL FIX. NOSE LANDING GEAR UPLOCK PROX SENSOR HARNESS REPLACED AND LANDING GEAR EXTENSION / RETRACTIONS SYSTEM OPS CHECKED SERVICEABLE. AIRCRAFT RETURNED TO SERVICE.

AFTER GEAR UP SELECTION, AND ALREADY IN A RIGHT TURN, TO FOLLOW THE DEPARTURE ROUTE, WE EXPERIENCED AN ABNORMAL YAW TO THE RIGHT FOR 1 SECOND AND THEN BACK TO NORMAL. WE EXPECTED A BIRDSTRIKE. NORMAL FLIGHT AND NORMAL LANDING. AFTER LANDING I PERFORMED AN IMMEDIATE INSPECTION AFTER DISEMBARKING OF THE PAX, AND FOUND THE RT ENGINE UPPER COWLING MISSING. PAX DIDN’T REALIZE ANYTHING. FURTHER FLIGHT WAS CANCELLED.

FLIGHT LT 3606 SCHEDULED FROM STR TO VIE DIVERTED TO MUC BECAUSE DURING CLIMB AT FL 260 ENGINE NR 2 OIL PRESS WARNING CAME ON SEVERAL TIMES. AFTER LEVEL OFF ENG NR 2 WAS REDUCED TO IDLE. OIL PRESSURE BETWEEN 30 AND 35 PSI. WITH INCREASING THRUST PRESSURE STILL BETWEEN 30-35 PSI. ENGINE WAS KEPT IN IDLE FOR THE REMAINDER OF THE FLIGHT. OIL TEMPERATURE WAS DECREASING TO 48 °C. MAINTENANCE INSPECTED ENGINE INSPECTED. ENGINE COVERED WITH OIL. FOUND 2 LOOSE COUPLINGS ON A-SUMP PUMP. COUPLINGS TIGHTENED. FOUND DURING RUN UP OIL ACCUMULATION AROUND OIL LEVEL SENSOR ON OIL TANK. LEVEL SENSOR AND PACKINGS REPLACED, NO HELP. OIL TANK PRESSURIZED. FOUND CRACK IN OIL TANK. OIL TANK REPLACED. AIRCRAFT RETURNED TO SERVICE.

UPON GEAR RETRACTION, GEAR WOULD INDICATE UP AND LOCKED FOR A FEW SECONDS, THEN THE AMBER HASHED INDICATIONS. THIS REPEATED SEVERAL TIMES ACCOMPANIED BY LOUD THUMPING NOISE. THERE WERE NO OTHER CAUTION, WARNING OR AURAL WARNINGS REPORTED. THE GEAR WAS EXTENDED NORMALLY AND ALL GEAR INDICATED DOWN AND LOCKED. AIRCRAFT RETURNED AND LANDED WITHOUT FURTHER INCIDENT. MX WAS ABLE TO DUPLICATE DEFECT WHILE AIRCRAFT WAS ON JACKS. MAIN LANDING GEAR VALVE DETERMINED TO BE AT FAULT. VALVE REPLACED AND OPS CHECKED SERVICEABLE. VALVE 750005000 0720 LEAKING MLG 15799/12794, AIRFRAME TIMES.

THE OPERATOR HAS REPORTED THE FOLLOWING, DURING TAKEOFF ON A FLIGHT THE NR 2 ENGINE UPPER ENGINE COWL DETACHED. THE CREW NOTED A SLIGHT BUFFET DURING THE EVENT AND CONTINUED TO THE DESTINATION AIRPORT WHERE IT WAS DISCOVERED THAT THE COWL WAS MISSING. THIS WAS THE FIRST FLIGHT AFTER THE AIRCRAFT HAD UNDERGONE AN A CHECK IN MAINTENANCE. THE CRJ TECHNICAL HELP DESK AND ISE ARE SUPPORTING THE OPERATOR. NOW THE ANSWERED QUESTIONS. WHICH SIDE IS THE COWLING IS MISSING. RT SIDE UPPER NOSE ACCESS COWLING , PROVIDE P/N OF THE MISSING COWLING, PN:228-50080-802 SN:SBRJNACC0 1183, ANY OTHER DAMAGES TO THE ACFT. NO DAMAGES, PILOT REPORTED AFTER DETAILED INSPECTION, CAN YOU EXPLAIN WHY THE COWLING DEPARTED THE ACFT. IT IS ASSUMED THAT AN ERROR APPEARED DURING A-CHECK, IT WAS THE FIRST FLIGHT AFTER THE A CHECK, IS SB 601R-71-007 HAS BEEN PERFORMED ON THE ACFT, THIS SB IS NOT APPLICABLE FOR THIS AIRCRAFT. IT SHOULD BE PERFORMED BEFORE DELIVERY. THERE WAS NO ENGINE CHANGE SINCE DELIVERY. ANY REPORT FROM THE PILOTS, HAVE THEY NOTICE BUFFETING OF ABNORMAL CONDITION. THE COWLING WAS LOST DURING TAKEOFF, AND IT IS STILL NOT FOUND ON THE AIRPORT AREA. THE PILOT NOTICED ONLY A LIGHT FORCE IN THE NOSE UP DIRECTION DURING TAKEOFF, LIKE A BIRD STRIKE. HE EXPLAINED THAT THERE WAS NO BUFFETING OR ANY OTHER FORM OF ABNORMAL CONDITION NOTICEABLE DURING THE WHOLE FLIGHT AFTER TAKEOFF. COWLING WAS REPLACED. TFH: 12205/THC: 10091.

WHILE RIGGING THE PARKING BRAKE CABLE SMALL BALLS WERE FOUND AT THE BOTTOM OF THE COMPARTMENT HOUSING THE BRAKE CONTROL VALVE PN 600-88101-125. UPON FURTHER INVESTIGATION IT WAS DISCOVERED THAT THE TWO FOLLOWER BEARINGS FOR THE PILOT BRAKE PEDAL INPUT TO THE BRAKE CONTROL VALVE WERE DESTROYED. PART CYCLES, TSN AND TSO ARE UNKNOWN ON THE BEARINGS.

(CAN) ACFT ON APPROACH, PILOT REPORTED ACFT HIT BY LIGHTNING IN THE NOSE AREA. IMMEDIATELY AFTER LIGHTNING STRIKE, RT ENG FLAMED OUT. PILOT DID NOT TRY TO RE-START RT ENG. ACFT LANDED, MX PRESENTLY INSPECTING THE ACFT. FDR TO BE DOWNLOADED. CUSTOMER CONTACTING MFG. ADDITIONAL INFORMATION 11 FEB 2009: FDR HAS BEEN DOWNLOADED. DOWNLOAD ON ITS WAY . SHOULD BE HERE TODAY AT NOON. ACFT AND ENGINE (EXTERIOR) INSPECTIONS CARRIED OUT. DAMAGES ARE MINOR. FUEL SAMPLES TAKEN. BOTH LT ENG AND RT ENG FUEL FILTERS ARE BEING DRAINED FOR QTY COMPARISON RT ENG CHIP DETECTOR OK. RT ENG BSI TO BE DONE ASAP. IT IS CONFIRMED THAT THE CONT IGN WAS ON PRIOR TO THE EVENT. DFDR CHECK SHARED WITH MFG. RT ENGINE DOES NOT FLAME OUT DURING PERIOD 180.22.23 THRU 180.23.10 BUT DOES ROLL BACK AND ITT INCREASES SIGNIFICANTLY INDICATIVE OF AN HPC STALL. THE ENGINE REMAINS IN STALL AND "ROLLS BACK" ITT INCREASES. AT APPROX 180.23.11 INDICATED FUEL FLOW DROPS TO ZERO, ITT DECLINES RAPIDLY WHICH IS CONSISTENT WITH PILOT ACTION FUEL SHUT OFF IN RESPONSE TO "ROLL BACK" AND LIKELY EICAS LOW OIL PRESSURE INDICATION LOW AND LOP AURAL WARNING. NOTE LIGHTNING APPEARS TO HAVE HIT THE FWD RT FUSELAGE AT THE LOWER PART OF THE SERVICE DOOR "SCUFF PLATE". EICAS INDICATIONS? "LOW OIL PRESS" AURAL AND INDICATION WHEN N2 GOT LOW ENOUGH? YES. YOU CAN SEE WHERE THE LIGHTNING STRUCK AND HOW CLOSE TO THE RT ENGINE INTAKE IT WAS. THE ACFT IS IN MX BUT MX REPORTS ENTRY POINT AROUND SVCE DOOR SCUFF PLATE (RT SIDE). AFTER INSP, MX REPORTS LIGHTNING ENTRY POINT.

(CAN) ACFT TOOK OFF FOR A FLIGHT TEST. THE TEMPERATURE WAS ABOUT -5 DEGREES CELSIUS, CLEAR SKIES AND WINDS FROM THE WEST BETWEEN 5 AND 10 KNOTS DURING THE INITIAL CLIMB, THE LANDING GEARS WERE RETRACTED AND HAD AN INDICATION OF 2 GREEN LIGHTS, HOWEVER THE NLG INDICATED "IN TRANSIT". THE NLG DOORS REMAINED OPENED. THE SYS INDICATED "GEAR DISAGREEMENT". PILOT DECIDED TO EXTEND THE GEARS, 3 GREEN LIGHTS, AND LANDED. MX CHECKED THE (PROXIMITY SENSOR ELECTRONIC-UNIT) PSEU FOR A FAULT CODE AND IT INDICATED A FAULTY (PROXIMITY SENSOR PS3GA) NOSE GEAR UPLOCK. RETRACTIONS OF THE NLG WERE MADE AND MX CONFIRMED THE UPLOCK MECHANISM WAS NOT FUNCTIONING CORRECTLY. FOUND THAT THE (EMERGENCY RELEASE LEVER) WHICH IS PART OF THE UPLOCK ASSY, WAS STIFF, PREVENTING THE TARGET (WHICH IS PART OF THE “MANUAL RELEASE LEVER”) FROM GETTING IN PROPER RANGE WITH THE PROXIMITY SENSOR. IT WOULD APPEAR AS IF THOSE TWO PARTS HAD BEEN ASSEMBLED TOO TIGHTLY TOGETHER. THE (UPLOCK) MECHANISM WAS REPLACED IAW AMM. THE NLG WAS RETRACTED SEVERAL TIMES AND FOUND SERVICEABLE. P/N OFF: 16600-101, S/N OFF: DCL263-96 P/N ON: 1660-103, S/N: NLG/0656/01.

RIGHT ENGINE N1 VIB ICON AND THE PILOTS REPORTED THE FAN VIBES REACHING 2.9 MILS. UPON LANDING AT AN OUTSTATION, A MECHANIC WAS SENT TO INVESTIGATE THE ISSUE. AFTER REMOVING THE FORWARD SPINNER IT WAS FOUND THAT THE FORWARD "BLIND" CAP OF THE FAN SHAFT HAD COME LOOSE AND WAS BANGING AROUND THE AFT SPINNER CAUSING THE IMBALANCE. DAMAGE WAS FOUND ON THE FORWARD BLIND CAP, RETAINING RING, AND AFT SPINNER. THE ENGINE FLEW EXACTLY 6 CYCLES BEFORE THE BLIND CAP LIBERATED FROM THE FAN SHAFT. BLIND CAP MOVEMENT IS BELIEVED TO BE DUE TO IMPROPER SEATING OF THE RETAINING RING DURING INSTALLATION, WHICH MIGHT BE A CONSEQUENCE OF RETAINING RING DEFORMATION.

"PILOT REPORTED A SNAG "RT ENGINE N1 READING 86 PERCENT, WHILE TARGET N1 WAS 88 PERCENT. RT ENGINE FAILED TO ACHIEVE TARGET N1 AS PER FCOM". POWER ASSURANCE CHECK WAS CARRIED OUT BY THE ENGINEER AND THE CHECK FAILED. THE DATA OF THE POWER ASSURANCE CHECK IS PROVIDED IN THE ATTACHMENT TITLED "ENGINE DATA". ONCE THE POWER ASSURANCE CHECK FAILED, I ASKED THE OPERATOR WHEN WAS THE LAST TIME AN ENGINE PERFORMANCE RECOVERY WASH WAS DONE. THEY SAID THAT IT WAS MORE THAN AN YEAR. I ASKED TO DO A RECOVERY WASH ON THE ENGINE AND AGAIN DO THE POWER ASSURANCE CHECK. THE DATA FOR THE POWER ASSURANCE CHECK DONE AFTER THE WASH ARE GIVEN IN THE ATTACHMENT TITLED "ENGINE DATA 2". THE CHECK FAILED EVEN THEN. THE AIRCRAFT IS FITTED WITH CF348C5-B1 ENGINES. PROPULSION FROM ENGINE GRD RUNS CRJ ISE RECOMMENDED TO CHECK ENGINE FMU MAIN FUEL FILTER AND OTHER CHECKS. N1 CONFIRMED TO BE LOW BUT INDICATED FUEL FLOW ABNORMALLY HIGH. CONFIRMED FUEL FILTER HAD "RUPTURED" AND INDICATION OF FILTER "SPINNING". LOW POWER (N1) AND HIGH FUEL FLOW CONSISTENT WITH FILTER RUPTURE AND AFFECT ON FMU MAIN FUEL METERING VALVE AND METERING VALVE POSITION WHICH IS USED BY FADEC TO COMPUTE AND INDICATE FUEL FLOW IN PPH TO EICAS. GE REFUSED HELP AND SUPPORT TO BOMBARDIER OR ALLIANCE REF "NON GE NON OEM PART INVOLVED". MAIN FUEL FILTERS ARE COMMON PN`S FOR ALL CRJ 100-200 AND 700-900. MAIN ENGINE FUEL FILTER FAA PMA PN 7582591 MFG PMA HOLDER PTI TECHNOLOGIES OXNARD CA USA N/K RUPTURED ALSO "SPINNING" OF FILTER NOTED TSN: 3472 HRS, CSN: 2763 CYCLES.

(CAN) LANDING GEAR DISAGREE WARNING AFTER GEAR RETRACTION IN THE CLIMB. ALL THREE GEAR INDICATED GREEN WHEN HANDLE WAS IN THE UP POSITION. REMOVED AND REPLACED PSEU IAE MM 32-61-01. C/W POWER UP TEST OF LANDING GEAR PSS IAW MM 32-61-00. C/W LANDING GEAR OPS TEST IAW MM 32-30-00. OPS CHECK GOOD. A REQUEST OF THE PSEU NVM AS BEEN MADE. MUST WAIT UNTIL UNIT IS RECEIVED BY VENDOR TO PERFORM DOWNLOAD. PSEU 8-953-04 D280 10826:08/7737/9068:50

PILOT REPORTED LEFT AIR CONDITIONING PACK BLOWS HOT AIR REGARDLESS OF TEMPERATURE SELECTION. SYSTEM INTERROGATED AND DETERMINED THE LOW PRESSURE MODULATING AND SHUTOFF VALVE DEFECTIVE. CLOSE INSPECTION OF VALVE PLENUM DETERMINED BUTTERFLY VALVE HAD BEEN EJECTED DOWNSTREAM OF PLUMBING AND WAS RETRIEVED WITH ALL HARDWARE INTACT. THE SHAFT OF THE VALVE THAT CONNECTS THE BUTTERFLY VALVE HAD BROKEN, RENDERING THE VALVE INOPERATIVE REGARDLESS OF TEMPERATURE INPUT. CHALLENGER 604 IPC REFERENCE 21-51-01 FIGURE 1 ITEM NR 10.

(CAN) WHILE IN SCOTTSDALE, ARIZONA, DURING THE PREFLIGHT CHECK THE CREW NOTICED THE APU INLET WAS DENTED AND APU INLET BELLOWS WAS RIPPED. NO OPERATIONAL ABNORMALITY WERE NOTICED BY THE CREW PRIOR TO THE DISCOVERY. MAINTENANCE TECHNICIAN WAS DISPATCHED AND BOTH PARTS WERE REPLACED WITH NEW AND APU INLET CHECKED FOR HIDDEN DAMAGE AS PER CL605 MM CHAPTER 49-14. BOMBARDIER IS INVESTIGATING THE POSSIBLE CAUSE. I HAVE LISTED THE MANUFACTURER AS CANADAIR BUT THE CORRECT MANUFACTURER NAME IS SENIOR AEROSPACE WHICH IS NOT LISTED ON THIS PROGRAM.

DURING THE REPLACEMENT OF THE BEARING IN THE ELEVATOR BELL CRANK ASSY, THE LINKS AND SPACERS WERE THE ELEVATOR CABLE ATTACHES WERE FOUND WORN, SPACERS WERE WORN RIGHT THROUGH (PN C2CF349ND). THE AN3 BOLT THAT ATTACHES LINK IN PLACE WAS WORN 1/3 THROUGH. SPACERS AND BOLTS WERE REPLACED.

THEENGINE SUFFERED A COMPLETE LOSS OF POWER DUE TO A FAILURE OF A FIRST STAGE REDUCTION GEARBOX PLANETARY GEAR. FAILURE OF THE PLANETARY GEAR WAS THE RESULT OF INTERGRANULAR AND FATIGUE CRACKING TO THE PLANETARY GEAR BORE. THE BEARING SLEEVE SPUN WITHIN THE BORE DUE TO CONTAMINANTS WITHIN THE LUBRICATING OIL ENTERING THE BEARING SLEEVE AND PLANETARY GEAR SHAFT INTERFACE. CONTAMINANTS WITHIN THE LUBRICATING OIL WERE AS A RESULT OF NORMAL OIL CONTAMINANTS BEING ALLOWED TO BYPASS THE FINGER SCREEN OF THE PLANETARY GEAR. UPON DISASSEMBLY OF THE POWER SECTION IT WAS FOUND THAT THE FINGER SCREEN WAS DEFORMED AND RESULTED IN FRETTING OF THE MAIN O-RING ON THE FINGER SCREEN, AND FRETTING OF THE FINGER SCREEN BODY AND FINGER SCREEN CHAMBER. THE FRETTING GENERATED CONTAMINANTS THAT WERE ALLOWED TO BYPASS THE FINGER SCREEN AND WEAR ON THE O-RING ALSO ALLOWED CONTAMINANTS TO BYPASS THE SCREEN. THE TEFLON PACKING WAS IMPROPERLY POSITIONED DURING THE LAST OVERHAUL OF THE ENGINE. THE IMPROPERLY INSTALLED TEFLON PACKING IMPOSED ABNORMAL STRESSES TO THE FINGER SCREEN AND RESULTED IN THE DEFORMATION OF THE FINGER SCREEN BODY. THE UNFILTERED OIL CAUSED HIGH TEMPERATURES TO THE PLANETARY GEAR RESULTING IN FAILURE OF THE GEAR AND ULTIMATELY FAILURE OF THE POWER SECTION OF THE ENGINE.

(CAN) DURING THE 12 MONTH INSPECTION OF THE MLG WHICH CONSISTS OF INSPECTING THE WELDS FOR CRACKS, A CRACK WAS DETECTED ON THE BOTTOM OF THE AFT ATTACH TUBE APPROX 5 INCHES FROM THE AFT GEAR ATTACH FITTING. (TC NR 20080924010)

ON APPLICATION OF BRAKES AFTER LANDING, THE LEFT PEDAL WENT TO THE MAX POSITION WITH NO EFFECT. LT BRAKE CABLE WAS FOUND SEPARATED APPROX 8 INCHES FROM THE VALVE IN THE UPPER FUSELAGE CLOSE TO A SMALL DIAMETER PULLEY. THIS CABLE WAS THE ORIGINAL INSTALLED IN 1981 AT THE FACTORY.

(CAN) DURING APPROACH, ONE ENG EXPERIENCED (2) UNCOMMANDED POWER ACCELERATIONS. THE CREW REDUCED THE POWER (PLA) AND SHUTDOWN ENGINE. THE ACFT MADE A SINGLE ENGINE LANDING AT PLANNED DESTINATION. THE FUEL CONTROL, FUEL PUMP AND EEC WERE REPLACED AND THE AIRCRAFT RETURNED TO SERVICE. MFG WILL CONTINUE INVESTIGATING THE EVENT AND ADVISE OF ROOT CAUSE ONCE ESTABLISHED.

(CAN) DURING T/O ROLL, THE CREW RECEIVED A FIRE WARNING ON NR 1 ENGINE AND IMMEDIATELY ABORTED T/O. IT IS UNKNOWN IF FIRE BOTTLES WERE DISCHARGED. THE A/C RETURNED TO THE RAMP WHERE INS REVEALED A FRACTURED 6 AND 7 BEARING OIL LINE, FRACTURED 1ST AND 2ND STAGE PT BLADES. HEAT RELATED DAMAGE WAS ALSO FOUND ON ACCESS PANNELS. THE LOCAL AUTHORITIES ARE INVESTIGATING THIS EVENT. MFG HAS OFFERED ASSISTANCE AND WILL FOLLOW UP.

WHILE CARRYING OUT LANDING GEAR RETRACTIONS, IT WAS FOUND THAT THE NOSE LANDING (NLG) WAS HUNG UP AT THE EXTENSION CYCLE. FURTHER INVESTIGATION REVEALED THAT THE NOSE LANDING (NLG) WAS NOT CENTERED AND THIS WAS PROBABLY CAUSED BY THE CENTERING SPRING NOT OVERCOMING THE BINDING FORCE. THIS RESULTED IN THE STEERING ACTUATOR BEING TWISTED UPWARDS AS A RESULT OF THE ADJUSTABLE LINK THAT WAS FOUND TO BE BROKEN. NOSE LANDING GEAR (NLG) ADJUSTABLE LINK ASSEMBLY REPLACED, FUNCTION CHECKED AND FOUND TO BE SERVICEABLE.

SHORTLY AFTER TAKEOFF, NR 1 HYD SYS WAS LOST, AIR CREEBEC FLIGHT 921 DECLARED AN EMERGENCY AND TURN BACK TO YUL AIRPORT ACFT LANDED SAFELY AND WAS TOWED TO GATE FOR PASSENGER OFFLOAD. AFTER MAINTENANCE INSPECTION THEY FOUND THE NR 1 ENG DRIVEN PUMP CASE DRAIN LINE RUPTURE AT THE FILTER AREA IN THE LT WHEEL WELL LINE AND PUMP WAS REPLACED AND THE ACFT WAS BACK IN SERVICE AROUND 11:00 AM.

SHORTLY AFTER TAKEOFF, THE NR 1 ENG HYD PUMP CAUTION LIGHT ILLUMINATED AND NR 1 HYD QTY GAGE WENT TO ZERO. THE CREW DECLARED AN EMERGENCY AND RETURNED FOR A SAFE FLAPLESS LANDING. MAINTENANCE FOUND IN THE LEFT WHEEL WELL A BROKEN HYDRAULIC LINE. THE LINE IS THE RETURN FROM THE ENGINE DRIVEN HYDRAULIC PUMP TO THE CASE DRAIN FILTER.

PIREP, STANDBY ATTITUDE INDICATOR FAILED. MADE GRINDING NOISE THEN TOPPLED. WOULD NOT RE-ERECT. LIGHTS REMAINED POWERED AND FLAG NOT IN VIEW. STANDBY INDICATOR REPLACED AND FUNCTION TEST SERVICEABLE. STRIP REPORT FROM AVIONICS SHOP ON SN 12395 STBY ATT IND, "FOUND THE ROTOR BEARINGS BOUND UP AND FUSE BLOWN. ERECTION AND DRIVE BEARINGS WORN." UNIT HAD 1196.5 HRS ON IT SINCE LAST REPAIR.

(CAN) DURING COMPLETION OF A-1 INSP TASK NR3220/51, EXTERNAL GENERAL VISUAL INSP OF THE NLG SHOCK STRUT AND DRAG STRUT, THE ENGINEER DISCOVERED A 9 INCH LONG CRACK ON THE AFT SIDE OF THE SHOCK STRUT CYLINDER. THE CRACK IS LOCATED BETWEEN THE SHOCK STRUT CYLINDER UPPER FLANGE AND THE CASTOR NUT LOCK PIN HOLE. THE NLG ASSY IS BEING REPLACED. THE NLG ASSY COMPLETE S/N IS DCL295/90/00R/00. THE SHOCK STRUT CYLINDER S/N IS DCL-796.

PROBLEM WAS FOUND TRYING TO START ACFT, AUX PUMP DOESN'T SUPPLY ENOUGH FUEL PSI TO OPEN MANIFOLD TO PRIME ENGINE FOR STARTING. (2) SPEED AUX/EMERGENCY FUEL PUMP WILL NOT PROVIDE ENGINE WITH SUFFICENT FUEL FLOW OR FUEL PRESSURE TO RUN ENGINE IF ENGINE DRIVEN FUEL BECOMES INOP IN AN EMERGENCY CONDITION. AIRPLANE ENGINE WILL NOT CONTINUE TO OPERATE IN FLIGHT. MFG REP. SAYS THAT IS CORRECT AND IT IS MFG PROBLEM. (ENGINE IS FINE AIRFRAME PROBLEM) MFG REP SAYS THE ACFT IS CERTIFIED. THEY ARE NOT ABLE TO ANWSER VERY MANY QUESTIONS. PROBLEM IS HOW DID THIS BECOME CERTIFIED IT IS AN UNSAFE CONDITION.

PILOT REPORT, UPON APPLYING POWER ON THE FIFTH TOUCH-AND-GO, WHILE ROLLING ON RUNWAY, THE ENGINE STOPPED RUNNING. ATC WAS IMMEDIATELY ADVISED THE ENGINE HAD QUIT. AFTER A FEW SECONDS WAS ABLE TO PRIME AND RESTART THE ENGINE WITHOUT PROBLEM. ATC WAS ADVISED OF THIS AND TAXIED DOWN RUNWAY AND EXITED RUNWAY ONTO TAXIWAY. THE ENGINE DID NOT GIVE ANY FURTHER TROUBLES AND SHUT DOWN AT THE MAIN BASE WITHOUT INCIDENT. MAINTENANCE RESPONSE, ADJUSTED FUEL PRESSURES AND MIXTURE SETTINGS. AIRCRAFT RUN UP AND LEAK CHECKED. AIRCRAFT RETURNED TO SERVICE SUBJECT TO SATISFACTORY TEST FLIGHT.

ENGINE SHUT DOWN AFTER TOUCHDOWN. FUEL SET UP PROCEDURE WAS PERFORMED IN ACCORDANCE WITH TCM SID97-3E AND GROUND RUN SATISFACTORY WITH ALL FUEL AND POWER INDICATIONS NORMAL. PAST EXPERIENCE WITH THESE ENGINE TYPES HAVE REVEALED THAT THEY CAN BE SENSITIVE TO AMBIENT TEMPERATURE AND PRESSURE CHANGES. ANOTHER CONTRIBUTING FACTOR MAY BE THE CONDITION OF THE FUEL SYSTEM COMPONENTS AND RELATED FUNCTIONS. THE FUEL SET UP PROCEDURE USUALLY CORRECTS THIS TYPE OF CONDITION.

LT MAIN TIRE BLOWN ON LANDING. AIRCRAFT REMAINED ON RUNWAY UNTIL MAINTENANCE CREWS WERE ABLE TO REPLACE THE DEFECTIVE WHEEL ASSEMBLY WITH A SERVICEABLE UNIT AND ENABLE THE AIRCRAFT TO TAXI BACK TO THE HANGAR FOR FURTHER INSPECTION.

PILOT REPORT, ENGINE FAILED ON RUNWAY FOLLOWING FULL POWER APPLICATION AFTER DOING TOUCH AND GO. ENGINE WAS RESTARTED AFTER SEVERAL ATTEMPTS AND TAXIED OFF. MAINTENANCE RESPONSE, FUEL SET UP PROCEDURE WAS PERFORMED IN ACCORDANCE WITH TCM SID97-3E AND GROUND RUN SATISFACTORY WITH ALL FUEL AND POWER INDICATIONS NORMAL. EXPERIENCE WITH THESE ENGINE TYPES HAVE REVEALED THAT THEY CAN BE SENSITIVE TO AMBIENT TEMPERATURE AND PRESSURE CHANGES. ANOTHER CONTRIBUTING FACTOR IS THE CONDITION OF THE FUEL SYSTEM COMPONENTS AND RELATED FUNCTIONS. THE FUEL SET UP PROCEDURE USUALLY CORRECTS THIS TYPE OF CONDITION.

INSPECTION OF THE RUDDER PEDAL ASSEMBLY REVEALED EXCESSIVE WEAR ON THE S TUBE CAUSED BY WEAR FROM THE RUDDER CABLE. THE WEAR WAS DETECTED BY VISUALLY CHECKING THE OUTER SURFACE OF THE S TUBE FOR MATERIAL DEFORMATION IN THE AREA(S) WHERE THE CABLE CAUSES THE MOST FRICTION ON THE INSIDE OF THE TUBE. THE AIRCRAFT RUDDER CABLES WERE DUE FOR REPLACEMENT AS MANDATED BY THE MANUFACTURER AT EACH 3000 HOURS TTSN. THIS DEFECT HAS BEEN ADDRESSED IN PREVIOUS REPORTS FROM THIS COMPANY. BASED ON THE NATURE OF THIS DEFECT WE REPLACED ALL RUDDER PEDAL ASSEMBLIES ALONG WITH THE CABLES AND HAVE AMENDED OUR INTERNAL MAINTENANCE TRACKING TO INSPECT/REPLACE AS NECESSARY THE PEDAL ASSEMBLIES AT 2500 HOUR INTERVALS. THIS ISSUE CONTINUES TO BE MONITORED.

DURING THE AIRCRAFT PREFLIGHT INSPECTION, THE PILOT REPORTED THAT THE PILOTS SIDE RUDDER PEDAL CABLE WAS STUCK IN THE FULL FORWARD POSITION. CLOSER INSPECTION BY MAINTENANCE REVEALED THAT THE PILOTS SIDE LT RUDDER CABLE HAD WORN THROUGH THE S TUBE THAT FORMS PART OF THE PEDAL ASSEMBLY. THE CABLE WAS EXPOSED THROUGH THE UPPER BEND OF THE S TUBE TO A POINT WHERE IT WAS JAMMED IN ITS WEAR HOLE THUS RESTRICTING THE SLIDE ACTION OF THE PEDAL ASSEMBLY ON THE ADJUSTMENT SLIDE TUBE. THE DEFECTIVE PART WAS REPLACED ALONG WITH THE OTHER PEDALS DUE TO THEIR SUSPECTED CONDITION NEARING THAT OF THE FAILED PART.

PILOT REPORTED THAT THE FLAPS WOULD NOT GO DOWN. INSPECTION REVEALED THAT THE ROLL PIN WHICH IS USED TO LOCK THE ROD END BEARING AT THE END OF THE FLAP ACTUATOR WAS MISSING ALLOWING THE FLAP ACTUATOR ROD TO ROTATE AND ALLOW FOR MIS-ALIGNMENT OF ALL FLAP CONTROL SWITCHES. REPAIRED FLAP ACTUATOR INSTALLED.

(CAN) PILOT REPORTED LOW OIL PRESSURE READING (15-20 PSI) WHILE PERFORMING TEST FLIGHT AFTER MAINTENANCE. NO LEAKS WERE DETECTED & ACFT GROUND RUN WITH AN EXTERNAL PRESSURE GAUGE TO VERIFY PRESSURE READINGS ON ACFT GAUGE. OIL PRESSURE RELIEF SPRINGS & PLUNGER REPLACED WITH NEW. OIL PRESSURE INDICATIONS REMAIN SUSPECT WHEN COMPARED TO INFORMATION PROVIDED IN ACFT FLT MANUAL. SECTION 2.5 (TR-1) OF AFM STATES THAT THE OIL PRESSURE SHOULD BE 30-60 PSI IN GREEN NORMAL OPS RANGE FOR TYPE. NORMAL OPERATING RANGE IS STATED AS 700-2800 RPM. BASED ON THIS INFORMATION, OIL PRESSURE SHOULD BE AT LEAST 30 PSI EVEN AT 700 RPM. PREVIOUS EXPERIENCE ON THIS ACFT TYPE HAS PROVEN THAT IT IS NOT UNCOMMON FOR OIL PRESSURE READING TO INDICATE BELOW 30 PSI AT VARIOUS PHASES OF FLT. ENGINE OPS MANUAL SECTION 6-22 STATES "IF OIL PRESSURE DROPS BELOW 30 PSI, AN ENGINE FAILURE SHOULD BE ANTICIPATED." HAVE CONSULTED WITH THE ENGINE MANUFACTURER AND AIRFRAME MANUFACTURER FOR CLARITY ON INTERPRETING THESE OIL PRESSURE INDICATIONS. VARIOUS IDEAS AND TROUBLESHOOTING TIPS HAVE BEEN EXCHANGED AND DOCUMENTED WITH AMO 62-92.

WHILE ATTEMPTING TO START THE ENGINE ON THE GROUND THE ENGINE TURNED OVER SEVERAL TIMES BUT THEN STOPPED. THE STARTER WAS STILL RUNNING AND COULD BE HEARD FROM THE COCKPIT. INVESTIGATION REVEALED THE STARTER BENDIX/GEAR HSG HAD CRACKED AND SEPARATED FROM THE CASTING.

THE NLG TUBULAR PIVOT ASSEMBLY SHOWS SIGNS OF FAILURE INDICATED BY THE MISALIGNMENT OF THE GUDGEON (D60-3233-82-31) WITHIN THE TUBULAR PIVOT SHAFT (D60-3233-82-32). THE MISALIGNMENT ALWAYS OCCURS ON THE LT GUDGEON AND ALWAYS OPENS TOWARD THE AFT DIRECTION WHEN THE GEAR IS DOWN AND LOCKED. FORCES ON THIS POINT ARE INDUCED BY THE NLG SPRING ASSEMBLY AND THE NLG ACTUATOR OVERDRIVING THE DOWN LOCK. REPEATED FORCES ON THE GUDGEON RESULT IN A BELLING DEFORMATION ON THE TUBULAR PIVOT SHAFT.

THE NLG TUBULAR PIVOT ASSEMBLY SHOWS SIGNS OF FAILURE INDICATED BY THE MISALIGNMENT OF THE GUDGEON (D60-3233-82-31) WITHIN THE TUBULAR PIVOT SHAFT (D60-3233-82-32). THE MISALINGMENT ALWAYS OCCURS ON THE LT GUDGEON AND ALWAYS OPENS TOWARD THE AFT DIRECTION WHEN THE GEAR IS DOWN AND LOCKED. FORCES ON THIS POINT ARE INDUCED BY THE NLG SPRING ASSEMBLY AND THE NLG ACTUATOR OVERDRIVING THE DOWN LOCK. REPEATED FORCES ON THE GRUDGEON RESULT IN A BELLING DEFORMATION ON THE TUBULAR PIVOT SHAFT.

THE NLG TUBULAR PIVOT ASSEMBLY SHOWS SIGNS OF FAILURE INDICATED BY THE MISALIGNMENT OF THE GUDGEON (D60-3233-82-31) WITHIN THE TUBULAR PIVOT SHAFT (D60-3233-82-32). THE MISALIGNMENT ALWAYS OCCURS ON THE LT GUDGEON AND ALWAYS OPENS TOWARD THE AFT DIRECTION WHEN THE GEAR IS DOWN AND LOCKED. FORCES ON THIS POINT ARE INDUCED BY THE NLG SPRING ASSEMBLY AND THE NLG ACTUATOR OVERDRIVING THE DOWN LOCK. REPEATED FORCES ON THE GRUDGEON RESULT IN A BELLING DEFORMATION ON THE TUBULAR PIVOT SHAFT.

THE NLG PIVOT ASSEMBLY FAILED BY SHEARING OF THE LT GUDGEON WITHIN THE TUBULAR PIVOT SHAFT. THIS FAILURE WAS NOTED UPON INSPECTION DUE TO A NLG INDICATION LIGHT NOT BEING ILLUMINATED DURING TAXI OUT. UPON RETURNING TO HANGAR THE FAILURE WAS NOTED. ALSO DISCOVERED WAS COMPOSITE DAMAGE ON THE LT WALL OF THE WHEEL WELL INFLICTED BY THE NLG COMPONENT DISLOCATION.

THE NLG TUBULAR PIVOT ASSEMBLY SHOWS SIGNS OF FAILURE INDICATED BY THE MISALIGNMENT OF THE GUDGEON (D60-3233-82-31) WITHIN THE TUBULAR PIVOT SHAFT (D60-3233-82-32). THE MISALIGNMENT ALWAYS OCCURS ON THE LT GUDGEON AND ALWAYS OPENS TOWARD THE AFT DIRECTION WHEN THE GEAR IS DOWN AND LOCKED. FORCES ON THIS POINT ARE INDUCED BY THE NLG SPRING ASSEMBLY AND THE NLG ACTUATOR OVERDRIVING THE DOWN LOCK. REPEATED FORCES ON THE GRUDGEON RESULT IN A BELLING DEFORMATION ON THE TUBULAR PIVOT SHAFT.

THE NLG TUBULAR PIVOT ASSEMBLY SHOWS SIGNS OF FAILURE INDICATED BY THE MISALIGNMENT OF THE GUDGEON (D60-3233-82-31) WITHIN THE TUBULAR PIVOT SHAFT (D60-3233-82-32). THE MISALIGNMENT ALWAYS OCCURS ON THE LT GUDGEON AND ALWAYS OPENS TOWARD THE AFT DIRECTION WHEN THE GEAR IS DOWN AND LOCKED. FORCES ON THIS POINT ARE INDUCED BY THE NLG SPRING ASSEMBLY AND THE NLG ACTUATOR OVERDRIVING THE DOWN LOCK. REPEATED FORCES ON THE GUDGEON RESULT IN A BELLING DEFORMATION ON THE TUBULAR PIVOT SHAFT.

DURING A PREFLIGHT INSPECTION, THE NLG DRAG BRACE PIVOT WAS FOUND TO BE LOOSE, BENT AND DAMAGED. THE FOLLOWING PARTS NEEDED TO BE CHANGED BECAUSE OF DAMAGE. 1 CAMP BOLT TUBE, 1 BRACE BEARING, 1 NIG BRACKET, 2 GUDGEON AND 2 BEARINGS. INVESTIGATION REVEALED THROUGH AN OPERATION TEST OF THE LANDING GEAR THAT THE NLG ACTUATOR IS PUTTING TO MUCH PRESSURE ON THE CAMP BOLT TUBE WHEN IN THE EXTENDED POSITION. THIS CAUSES THE CAMP BOLT TUBE TO BEND AND OVER TIME CAUSES THE ENDS TO FLARE OUT WHERE THE GUDGEONS AND INSTALLED. IF ALLOWED TO GO ON THE GUDGEONS COULD FALL OUT.

DURING IMC FLIGHT, THE RT ENGINE EXPERIENCED A RAPID 50 PERCENT POWER EXCURSION DOWNWARD TO 25 PERCENT AND RAPID SURGING THEREAFTER. AN EMERGENCY WAS DECLARED TO ATC, THE FLIGHT COMPLETED SAFELY. INVESTIGATION: ON THE FADEC ENGINE WIRING LOOM, THE PROPELLER CONTROL VALVE CONNECTOR CONTACTS WERE FOUND BURNED FROM ARCING, AND THE CONTACTS WERE MECHANICALLY SPREAD AND NO LONGER ABLE TO MAKE GOOD CONTACT WITH THE PROP CONTROL VALVE SPADE TYPE PINS. IF THIS CONNECTOR LOSES CONDUCTIVITY WITH THE PROPELLER CONTROL VALVE IN FLIGHT, THE PROP "FAIL SAFES" TO FEATHER, SHUTTING THE ENGINE DOWN. THIS OPERATOR HAS HAD SEVERAL PROBLEMS WITH THESES ENGINES DEEP SURGING, INCLUDING ONE IN-FLIGHT. (K)

LAST YEAR ABOUT THIS TIME WE WERE PERFORMING PERFORMANCE MODS ON THE ACFT. FOUND THE MAIN WING ATTACH POINTS TO HAVE EXCESSIVE PLAY ON BOTH WINGS WHEN ACFT WAS ON JACK. ENGINEERING LOOKED AT THE DATA WE PRESENTED TO THEM AND SAID "THE PLAY WAS WITHIN LIMITS." HAVE WORKED MANY ACFT AND NEVER SEEN THIS CONDITION ACCEPTABLE. IT HAS APPROX .020 TO .030 PLAY AT MAIN WING ATTACH FITTING. WHICH AT THE WING TIPS = OUT TO ABOUT 8 INCH. ENGINEERING DID NOT SPECIFY ANY ADDITIONAL INSP REQUIREMENTS. BELIEVE THIS WAS A COVER UP BY UPPER MANAGEMENT AT THE TIME SO THAT THE PRODUCT WOULD NOT BE DAMAGED. I AM HIGHLY CONCERNED THAT IS AN UNSAFE CONDITION AND NEEDS TO BE LOOKED INTO. THERE ARE MANY OTHER ITEMS ON THIS ACFT THAT NEED TO BE LOOKED INTO AS WELL SUCH AS AVIONICS SYS. FIRE PROTECTION AGENT IS SEVERALLY CORROSIVE WITH IN A DAY IF THE AGENT IS NOT FOUND LEAKING IT EATS STAINLESS STEEL AND ALUMINUM VERY QUICKLY. THE RT ENGINE IS EFFECTED SEVERAL BECAUSE THE IMPENDING BYPASS SWITCH FOR THE FUEL CONTROLLER IS DRIPPED ON BY THE AGENT AND COULD CAUSE AN UNSAFE CONDITION IF THERE IS BLOCKAGE WITH FUEL THE INDICATION MY NOT BE RECORDED. AT THE TIME,DID NOT REPORT IT BECAUSE MFG SCARED EMPLOYEES WITH THERE NONDISCLOSURE AGREEMENT AND TERMINATION IF ANY INFORMATION WAS LEAKED OUT. AC NR SERIAL NR 28 WAS THE AIRCRAFT THAT THE WINGS WERE FOUND LOSS ON.

DURING A POST FLIGHT INSPECTION, IT WAS NOTED THAT THE INNER PANE OF THE PILOTS WINDSHIELD HAD LOCALIZED HAZING IN THE CRITICAL VISION AREA, AS DEFINED BY EA500 AMM 06-117751, CHAPTER 56-00-00, FIG 601; JUST BEHIND THE GLARE SHIELD MOUNTED OPTIONAL ELECTRIC ATTITUDE INDICATOR. ACFT WAS FERRIED TO SERVICE CENTER, UNDER LIMITED ALTITUDE OF NO MORE THAN 12,000 (UNPRESSURIZED) FEET WITH AIRSPEED OF NO MORE THAN 250 KNOTS, FOR THE REPLACEMENT OF PILOTS WINDSHIELD. CONTROL NO. CUXA2009-00001

TRANSCRIPTION OF EMAIL RECEIVED FROM MRO: BRAKE PN 90000583-1PR SN MAR05-0550 AC PO C129RAER006 THIS BRAKE WAS RECEIVED IN PRETTY BAD CONDITION. IT LOOKS LIKE 2 ROTORS AND 1 STATOR FAILED DURING OPERATION. ATTACHED ARE SOME PICTURES OF THE BRAKE AND PIECES OF THE CARBON STACK THAT FELL OUT OF THE BRAKE. WE HAVE NOT DISASSEMBLED YET IN CASE SOMEONE NEEDS TO INVESTIGATE FURTHER. IT ALSO LOOKS LIKE SOME DAMAGE TO THE TORQUE TUBE WHERE WE CAN SEE IN THROUGH THE HOLE BY THE MISSING CARBON DISK.

TRANSCRIPTION FROM EMAIL RECEIVED FROM ENGINEERING DEPT, FAILED WITH 2048 LANDINGS SINCE INSTALLATION, MPN: 90000583-1PR WITH SN: APR05-0595 FOUND THE DAMAGE WAS FOUND AFTER REMOVAL OF THE WHEEL TO REPAIR FLAKING PAINT WHICH WAS IDENTIFIED IN THE 14TH OF JANUARY DURING A WTL M/W CHANGE AT THAT POSITION. NO DAMAGE TO THE BRAKE WAS NOTED AT THAT TIME. SO IN MERELY 7 DAYS WE HAD SUFFICIENT DETERIORATION OF THE CARBON CONDITION TO CAUSE FAILURE DESPITE HAVING DONE THE RECOMMENDED INSPECTIONS PER THE SB.

AIRCRAFT ON LOCAL FLIGHT TRAINING MISSION. INSTRUCTOR NOTED FLAPS CIRCUIT BREAKER POPPED WHEN FLAPS SELECTED UP. C/B RESET AS PER CHECKLIST, POPPED WHEN SELECTED UP A SECOND TIME. INSTRUCTOR ADVISED ATC OF FLAP DIFFICULTY AND RETURNED TO LAND WITH FLAPS IN LANDING, FULL DOWN, POSITION. ACFT WAS MET WITH AIRFIELD EMERGENCY RESPONSE, LANDED WITHOUT INCIDENT. MAINTENANCE BROUGHT ACFT INTO WARM HANGAR. FLAPS CYCLED REPEATEDLY WITHOUT FAULT. ACFT RELEASED SERVICEABLE. THE FLAP C/B EVENT IS TYPICAL OF TYPE IN COLD WEATHER (MID -20S C. AND COLDER). THE FLAP ACTUATOR IS AN `ACME` THREAD SCREW JACK DRIVEN BY FLEXIBLE DRIVES FROM THE FLAP MOTOR/TRANSMISSION ASSY. IT IS SUSPECTED THAT THE LUBRICANT IN THE ACTUATOR BECOMES TOO VISCOUS IN COLD WEATHER CAUSING EXCESSIVE RESISTANCE AT THE MOTOR. CONSEQUENTLY THE MOTOR IS SLOWED CAUSING THE HIGHER THAN NORMAL CURRENT DRAW, HENCE THE C/B POPS. OEM IS WORKING WITH OPERATOR TO DELIVER FLAP ACTUATORS WITH AN IMPROVED LUBRICANT THEY BELIEVE WILL IMPROVE THE COLD WEATHER RELIABILITY OF THE FLAP SYSTEM.

ON VISUAL INSPECTION OF THE NOSE LANDING GEAR DRAG BRACE AND BUNGEE SYSTEM, IT WAS FOUND THAT THE SHAFT LINKING THE DRAG BRACE TO THE BUNGEE WAS CRACKED AT THE BOLT HOLES. THE BOLT HOLES WERE FOUND TO ALSO BE ELONGATED. THE SHAFT WAS REMOVED AND THE HIDDEN PART OF THE SHAFT WAS SEVERELY CRACKED. SIX OTHER AERO COMMANDER 500B AND 500S ARE TO BE INSPECTED. TO DATE 2 SHAFTS ARE FOUND CRACKED.

AFTER TAKEOFF THE GEAR LANDING GEAR WAS SELECTED UP. ALL 3 "DOWN AND LOCKED" LIGHTS EXTINGUISHED BUT THE "IN TRANSIT" RED LIGHT REMAINED ILLUMINATED AND THE TOWER REPORTED THAT THE RIGHT MAIN GEAR WAS STILL EXTENDED. THE LANDING GEAR WAS SELECTED DOWN. THE NOSE AND LEFT MAIN GEAR "DOWN AND LOCKED" LIGHTS ILLUMINATED IMMEDIATELY BUT THE RED "IN TRANSIT" LIGHT REMAINED ILLUMINATED. AFTER APPROXIMATELY 30 SECONDS THE RIGHT MAIN GEAR "DOWN AND LOCKED" INDICATOR ILLUMINATED AND THE "IN TRANSIT" LIGHT EXTINGUISHED. THE AIRCRAFT LANDED AND RETURNED TO THE HANGAR. A VISUAL INSPECTION REVEALED NO OBVIOUS DAMAGE OR LEAKAGE. A REPLACEMENT ACTUATOR WAS INSTALLED AND THE AIRCRAFT RETURNED TO SERVICE. NO RECURRENCE HAS BEEN NOTED IN THE FOLLOWING 3 FLIGHTS DURING WHICH THE LANDING GEAR WAS RETRACTED AND EXTENDED NUMEROUS TIMES. THE FAILED ACTUATOR HAS BEEN RETURNED TO THE MANUFACTURER FOR INVESTIGATION.

OIL LEAKING FROM UPPER REAR CYLINDER HOLD DOWN STUD. THE OIL WAS FOUND TO RUN DOWN THE STUD AND EXIT ONTO THE CYLINDER FINS. OIL LEAK WAS ONLY NOTICED AT HIGHER TEMPS, AREA WOULD STAY CLEAN AT NORMAL RUN-UP CONDITIONS. THIS LEAK IS UNUSUAL AS THE STUD DOES NOT PROTRUDE INTO THE CASE (SO NO O-RING OR SEALANT IS REQUIRED). CASE WAS DISASSEMBLED AND NDT LOCATED THE CRACK ON THE INSIDE WEBBING OF THE CENTER CRANKSHAFT JOURNAL. CRACK WAS NOT VISUALLY DETECTABLE AND DID NOT CONTINUE TO THE OUTSIDE SURFACE OF THE CRANKCASE. CRANKCASE TO BE REPLACED WITH NEW.

OPERATING A SCHWEIZER 269 C MODEL SERIAL NUMBER S1575 TSN 6449.5 THE CONICAL BRGS WERE REPLACE WITH NEW AT 6338.8 AND A NEW BOLT WAS INSTALLED AT THIS TIME P/N 269A6092-3 WHICH WAS PURCHASE NEW FROM SCHWEIZER ON. THE RETORQUE WAS CARRIED OUT AT 6416.1 TSN NO MOVEMENT WAS NOTED AT THIS TIME. THE AIRCRAFT WAS IN FLIGHT AT THE TIME OF THE OCCURRENCE. THE PILOT WAS TURNING SHORT FINAL WHEN HE FELT A VIBRATION SUSPECTED ENGINE PROBLEMS. HE COMPLETED A RUN ON LANDING UNDER POWER WHILE MAINTAINING DIRECTIONAL CONTROL. UPON LANDING THE AIRCRAFT WAS NOT PERFORMING CORRECTLY AND A VIBRATION WAS FELT. THE PILOT SHUT DOWN AND MAINTENANCE WAS BROUGHT OUT TO AIRCRAFT. UPON INSPECTION OF AIRCRAFT IT WAS NOTED THE FLAPPING HINGE BOLT WAS SHEARED AND THE AIRCRAFT WAS THEN GROUND HANDLED TO THE HANGAR FOR FURTHER INSPECTION.

SPRAG ASSY FAILED 300 HR INSPECTION DUE TO 2 BROKEN CORNERS OF 2 DIFFERENT TEETH (SPRAGS). NO OTHER DAMAGE WAS FOUND TO THE ASSEMBLY AS A RESULT OF THIS INCIDENT. CLUTCH ASSY WAS WORKING FINE PRIOR TO INSPECTION. NEW SPRAG ASSY INSTALLED AND CLUTCH RETURNED TO SERVICE.

THE ACFT USES PRECISE FLIGHT SPEEDBRAKES. 31 AUG 07, FOUND A PROBLEM WITH THE SPEEDBRAKES AND WERE REPLACED. THE SPEEDBRAKES HAVE BEEN OVERHAULED (5) TIMES SINCE. 2 MAR 09, DEPLOYED THE SPEEDBRAKES IN FLIGHT AND THE ACFT YAWED TO THE RIGHT. APPARENTLY ONE SPEEDBRAKE DEPLOYED BEFORE THE OTHER. INSPECTED THE SPEEDBRAKES AND FOUND SEVERAL DISCREPANCIES. THE PROBLEM COULD BE A VERY TIGHT FIT OF THE RT SPEEDBRAKE. SCRAPED AND SANDED THE RT SPEEDBRAKE HOLE. THE SPEEDBRAKES ARE CURRENTLY BEING OVERHAULED. THIS IS A SAFETY OF FLIGHT PROBLEM. WHEN THE POWER WAS REMOVED THE RT SPEEDBRAKE AND WOULD STAY DEPLOYED. (K)

WHILE CLIMBING AT 17,000 FT, APPROX 40 MINUTES INTO FLIGHT, HEARD LOUD BANGS COMING FROM THE RT ENG. THE ENGINE SHUTDOWN, AND THE A/C LANDED (UNSCHEDULED) WITHOUT INCIDENT. UPON VISUAL INSPECTION, IT WAS FOUND THERE WAS CONTAINED TURBINE FAILURE, AND BLADES WERE MISSING. THE ENGINE IS BEING REMOVED FOR TEARDOWN ANALYSIS, AND INVESTIGATION BY MAINT.

(CAN) FLAP POWER DRIVE UNIT FOUND LEAKING FROM SOLENOID VALVE, THIS IS THE (4TH) UNIT WE HAVE HAD THE SAME FAULT ON, NO SDR WAS SUBMIITTED ON THE PREVIOUS UNITS. THE FIRST UNIT SN 1385P HAD 1217.9 TT WHEN IT STARTED TO LEAK FROM SOLENOID, REPLACED 2008/04/11, NEXT UNIT SN 1101 WAS ON FOR 29.1 HOURS, REPLACED 2008/05/11, NEXT UNIT SN 1387 WAS ON FOR 99.6 HOURS REPLACE 2009/07/11, UNIT SN 1322 WAS INSTALLED FOR 279.8 HOURS. ACCORDING TO MM NO LEAK RATE FROM THE SOLENOID VALVE IS GIVEN.

(CAN) AFTER DEPARTING, THE ACFT SUFFERED A CENTER WINDSHIELD OUTER PANE FAILURE. THE CREW REDUCED CABIN PRESSURE AND THE ACFT RETURNED TO POINT OF DEPARTURE AND LANDED WITHOUT FURTHER PROBLEM. MX REPLACED THE CTR WINDSHIELD AND THE ACFT WAS RETURNED TO SERVICE.

DEPARTING YZF THE CREW OBSERVED A LOW OIL INDICATION ON NR 3 PROPELLER IN THE CLIMB WITH NORMAL RPM. THE PROPELLER RPM INCREASED TO 103 PERCENT SHORTLY AFTER AND THE CREW SHUT THE NR 3 ENGINE DOWN. THE AIRCRAFT RETURNED TO POINT OF DEPARTURE AND LANDED WITHOUT FURTHER PROBLEM. MAINTENANCE REPLACED THE NR 3 PROPELLER, AND PROP VALVE HOUSING AND THE AIRCRAFT WAS RETURNED TO SERVICE. MAINTENANCE OBSERVED SOME DAMAGE ON 3 PROPELLER BLADE SEALS AND IT IS UNCONFIRMED THAT THE FLUID LOSS INDUCED A VALVE HOUSING PROBLEM OR IF THE OVER SPEED CONDITION WAS RESULTANT FROM AN ACTUAL PROP VALVE FAILURE. TEARDOWN RESULTS ARE BEING REQUESTED FROM THE VENDOR.

DURING PHASE DEPOT INSP THE LWR FWD LT RING FITTING, PN 903220-1 WAS FOUND FRACTURED THRU THE HORIZONTAL BOTTOM FLANGE AT THE FILLET NUT PLATE HOLE AND HAD PROPAGATED ACROSS THE BOTTOM WEB AND UP THE VERTICAL THROUGH TWO FASTENERS; THE FIRST .1875 INCHAND THE SECOND .2500 INCH DIAMETER WHERE THE FRACTURE TERMINATED. THE TOTAL LENGTH IS APPROXIMATELY 2.6250 INCHES.

(CAN) DURING AN AVIONICS SYS MODIFICATION, DIMMING CONTROL RELAY PANEL P/N 8090986-401 WAS REMOVED FROM PILOT`S SIDE OF COCKPIT FOR ACCESS. THE INSULATION BAG, P/N 80-909156-11C WAS OBSERVED TO HAVE (2) TWO BURN HOLES IN IT, APPARENTLY CAUSED BY THE (2) RESISTORS THAT ARE ATTACHED TO THE DIMMING CONTROL RELAY PANEL. UPON FURTHER INVESTIGATION, IT APPEARS THAT THE INSULATION BAG (P/N 80-909156-11C) WAS NOT INSTALLED IN THE PROPER POSITION.

CREW REPORTS OF INTERMITTENT FLAP OVERSPEED WARNINGS WITH FLAPS UP. SOMETIMES ASSOCIATED WITH FLAP AND PUSHER WARNINGS. NO MAINTENANCE CODES WERE NOTED FROM FCWU. SLOTTED BUSHING IN FLAP DRIVE ARM ASSY WAS FOUND UNBONDED, UPON ROTATION, THE IMPROPER ORIENTATION OF THE SLOT WOULD CAUSE THE ARM TO HAVE EXCESSIVE PLAY. THIS IN TURN WOULD ALLOW THE SWITCH ACTUATOR ARM TO COME OFF THE UPLIMIT SWITCH INTERMITTENTLY IN FLIGHT AND REMOVE THE "FLAPS UP" SIGNAL TO THE FLAP CONTROL AND WARNING UNIT, THE STICK PUSHER COMPUTERS AND THE AIRSPEED WARNING SYSTEMS. BOTH LEFT AND RIGHT ARM ASSYS WERE REPLACED WITH NEW AS RT WAS FOUND WITH UNBONDED BUSHES.

ON MORNING GROUND RUN NR 2 GENERATOR WOULD NOT TAKE A LOAD MORE THAN 60 AMPS AS READ OFF EIS. NR 2 GENERATOR WOULD GO OFF AND ON LINE BY IT SELF (FLASHING WARNING) ON ANNUNCIATOR PANEL AND OVERHEAD PANEL. REPLACED GENERATOR WITH SERVICEABLE UNIT, DEFECT GONE.

DURING CLIMB AND CRUISE BOTH PRIMARY FLIGHT DISPLAYS (PFD) INDICATED A 12-20 DEGREE LT TURN WHILE THE ACFT WAS IN A WINGS LEVEL CONDITION, WHICH WAS CONFIRMED BY THE VISUAL METEOROLOGICAL CONDITIONS AND THE ATTITUDE INDICATION ON THE ELECTRONIC STANDBY INSTRUMENT SYSTEM (ESIS). THIS IS BEYOND THE 10 DEGREE ATTITUDE ROLL ATTITUDE OFFSET DESCRIBED IN AD 09-04-14 AND POH TEMPORARY REVISION 9. THIS DISCREPANCY OCCURRED AFTER MFG "BUILD 5" HAD BEEN ACCOMPLISHED ON THIS ACFT. BOTH CHANNELS OF THE ADAHRS UNIT TEST NORMALLY IAW AMM 12-B-34-25-00-00S-903S-S WITH NO DOWNLOADABLE FAULTS.

DURING AN ANNUAL INSPECTION THE LT MAIN WHEEL WAS REMOVED TO REPACK THE WHEEL BEARINGS WHEN IT WAS DISCOVERED THAT OTBD BRAKE DISC ON THE LT BRAKE WAS BROKEN INTO SEVERAL PIECES. THE BRAKE ASSY WAS REMOVED AND REPLACED.

ACFT WAS STARTED AND TAXIED TO ACTIVE RUNWAY FOR DEPARTURE. A NORMAL RUN-UP WAS PREFORMED AND NO PROBLEM WAS NOTED. ON TAKEOFF NORMAL RPM WAS NOTED AND AS THE ACFT BECAME AIRBORNE AND ON INITIAL CLIMB OUT ENGINE ROUGHNESS WAS NOTED AND THEN AS THE PILOT MADE A DECISION TO ENTER A DOWNWIND FOR A RETURN TO THE FIELD A NOTICEABLE ROUGHNESS AND POWER LOSS WAS NOTED AND AN UNEVENTFUL RETURN TO THE FIELD WAS MADE. UPON AN INVESTIGATION OF THE ENGINE COMPARTMENT AND OF THE INDUCTION SYSTEM IT WAS NOTED THAT THE CARBURETOR BOWL COULD BE MOVED. THE (4) FOUR SCREWS HOLDING THE BOWL ON HAD THE LOCKS ATTACHED AND THE SCREWS ARE LOOSE. THIS CAUSED AN UNEVEN FUEL FLOW AND DISRUPTION OF FLOW CAUSING ENGINE ROUGHNESS AND POWER LOSS. THE CARBURETOR HAD BEEN OVERHAULED BY THE COMPANY THAT HAD OVERHAULED THE ENGINE WITH ONLY 125.5 HR T.I.S. IN THIS CASE THE IMPROPER INSTALLATION OF BOWL TO THROTTLE BODY ASSY OR THE USE OF A BOWL ASSY. WITH DEFECTIVE THREADS WAS THE CAUSE OF THE DEFECT.

ON APPROACH TO YBL, THE PILOT DETECTED BURNING OIL AND SAW SMOKE, THE PILOT LANDED THE AIRCRAFT AND NOTICED A LARGE AMOUNT OF OIL ON THE BELLY OF THE AIRCRAFT AND DRIPPING FROM THE COWLS, AFTER THE TECHNICIAN INVESTIGATED THE OIL LEAK HE FOUND A HOLE WHERE A PLUG SHOULD BE IN THE ACCESSORY CASE, AFTER A BRIEF SEARCH FOR THE MISSING PART AND NOT FINDING THE PLUG THE TECH INSTALLED A NEW PLUG (PN AN913-3) AND LOCKWIRED.

THE AIRCRAFT LANDED WITH INDICATIONS OF A SOFT NOSE TIRE ON THE ROLL OUT. ON THE TAXI TO THE HANGAR, THE TIRE WENT COMPLETELY FLAT. THE TIRE WAS REPLACED AND THE AIRCRAFT CONTINUED ON IT`S SCHEDULED FLIGHT. WHEN THE TIRE WAS DISASSEMBLED, A "ROUND SHAPED" CUT WAS FOUND IN THE TUBE SIDEWALL. IT WAS THE SHAPE AND DIAMETER OF A AS3209-009 O-RING. AN O-RING OF THIS SIZE WAS FOUND BETWEEN THE TUBE AND THE TIRE WALL. IT IS BELIEVED THAT THE O-RING WAS TRAPPED INSIDE THE TIRE DURING BUILD UP AND THE MOVEMENT OF THE TUBE CAUSED THE CHAFING WHERE THE O-RING WAS TRAPPED. A REVIEW OF PROCEDURES HAS BEEN COMPLETED WITH THE MAINTENANCE SHOP TO AVOID THIS ISSUE AGAIN. IT HAS BEEN SUBMITTED AS AN SMS REPORT.

(CAN) WHEN THE ACFT TOUCHED DOWN, A VIBRATION WAS FELT IN THE NOSEGEAR. THE NOSE TIRE WAS COMPLETELY FLAT AFTER THE ACFT STOPPED. THE ACFT WAS TOWED TO THE HANGAR AND THE WHEEL REMOVED. THE VALVE STEM HAD SHEARED OFF AS WELL AS THERE BEING SOME DAMAGE TO THE TUBE. IT WAS BELIEVED THAT MOST OF THE DAMAGED WAS CREATED DURING THE TOWING OF THE ACFT. A NEW TIRE AND TUBE WERE INSTALLED ON THE ACFT.

(CAN) AFTER TAKEOFF THE RT BOOST PUMP CIRCUIT BREAKER POPPED. THE FLIGHT CREW THEN RETURNED TO BASE. UPON INVESTIGATION BY MX IT WAS FOUND THE RT BOOST PUMP HAD SEIZED. THE BOOST PUMP WAS REPLACED AND ACFT RETURNED TO SERVICE.

DURING A ROUTINE INSPECTION IT WAS DISCOVERED THAT THE TAB ON THE STARTER THAT THE BELT TENSIONER PULLEY ATTACHES TO WAS SHEARED AT THE HOUSING. THE PART ONLY HAD 128.4 HOURS ON IT AND WAS RETURNED FOR WARRANTY. THE PART MANUFACTURER IS SKY TEC BUT I COULDN`T ENTER THAT BRAND NAME.

(CAN) IN CRUISE, THE PILOT NOTICED THAT THE RT ENGINE OIL TEMP WAS CLIMBING TO THE RED LINE WITH A DECREASE IN OIL PRESSURE. THE PILOT DECIDED TO SHUT DOWN AND SECURE THE ENGINE ONCE THE OIL TEMP REACHED THE RED LINE. THE PILOT RETURNED TO BASE WITHOUT INCIDENT. MX FOUND THAT THE VERNATHERM VALVE SPRING WAS BROKEN AND WAS PROTRUDING OVER THE SEAT WHICH WOULD NOT ALLOW THE VERNATHERM TO DIRECT THE OIL THROUGH THE OIL COOLER. A NEW VERNATHERM VALVE WAS INSTALL AND THE ENGINE OIL TEMP RETURNED TO NORMAL.

PILOT REPORT ENGINE FUEL PUMP ON APPROACH AND VACUUM PUMP FAILURE. FUEL PUMP AND VACUUM PUMP FOUND BOTH WITH SHAFT BROKEN. BOTH REPLACED, ENGINE DIDN`T START. ENGINE REMOVED AND SENT FOR REPAIR. REPORT FROM REPAIR SHOP RECEIVED WITH INDICATION OF METAL CONTAMINATION AND SEVERE DAMAGED TO ACCESSORIES GEAR MAYBE CAUSE BY FAILURE OF FUEL PUMP. PUMP STILL IN OUR HAND WE WOULD GET AN EXPERTISE ON THE UNIT.

BOTH RT AND LT ALTERNATOR BELTS BROKE IN FLIGHT, POSSIBLE CAUSE WAS BOTH ENGINES MAY HAVE BEEN STARTED WITH NOSE COWL PLUG WOOL BLANKETS INSTALLED, THUS, BINDING UP THE PULLEYS AND/OR BELTS WHILE THE REST OF THE MOVING PARTS SEVERELY CUT INTO THE BELTS. BOTH BELTS HAD LESS THAN 10 HOURS TIME IN SERVICE. (K)

AFTER REMOVING THE AIRCRAFT FROM JACKS, I GRABBED ONTO THE RIGHT SIDE WING TIP TANK AND GAVE THE AIRCRAFT A SHAKE IN AN EFFORT TO SETTLE THE OLEO STRUTS. IN DOING SO I NOTICED AN OBVIOUS MOVEMENT BETWEEN THE TIP TANK AND WING STRUCTURE. I REMOVED THE FAIRINGS WHICH COVER THE MOUNTING POINTS FOR THE TIP TANK AND THEN GAVE IT ANOTHER SHAKE. AT THIS POINT IT WAS OBVIOUS THAT THE TIP TANK MOUNTING BOLTS WERE LOOSE. TO BE MORE SPECIFIC, THE TIP TANK IS HELD INTO PLACE BY 6 BOLTS. ONE 4/16 INCH BOLT AT THE FORWARD END, ONE 4/16 INCH BOLT AT THE AFT END AND FOUR 3/16 INCH BOLTS IN THE CENTER WHICH ATTACH TWO (APPROX 4 INCH SQUARE) MOUNTING PLATES. THE MOVEMENT WAS NOTED AS BEING BETWEEN THESE TWO PLATES IN THE CENTER. REMOVED ONE OF THE CENTER BOLTS TO INSPECT FOR WEAR IN THE BOLT HOLE, NO ELONGATION WAS NOTED, AND THE BOLTS WAS RE-INSTALLED. THE SURROUNDING STRUCTURE, BOTH ON THE WING AND ON THE TIP TANK, WAS INSPECTED WITH NO FAULTS FOUND. ALL OF THE MOUNTING BOLTS WERE TORQUED AS PER THE MANUFACTURERS MAINTENANCE MANUAL AND THE FAIRING WAS RE-INSTALLED. I INSPECTED THE LEFT WING TIP TANK AND THERE WAS NO MOVEMENT FOUND. REMOVED THE FAIRING AND TORQUE CHECKED ALL OF THE MOUNTING BOLTS, ALL OF THE CENTRE (FOUR) BOLTS WERE FOUND TO BE UNDER TORQUED, BUT THE FORWARD AND AFT BOLTS WERE TIGHT. I TORQUED THE CENTER BOLTS AND RE-INSTALLED THE FAIRINGS. WE HAVE SCHEDULED A 50 HOUR INSPECTION INTERVAL ON THESE MOUNTING BOLTS AND WILL RE-EVALUATE THE INTERVAL AS TIME GOES ON.

WHILE COMING IN FOR A LANDING, CUSTOMER WENT TO EXTEND LANDING GEAR AND GEAR WOULD NOT EXTEND. USED EMERGENCY GEAR EXTENSION AND LANDED ACFT WITH NO INCIDENT. UPON TROUBLESHOOTING FOUND RETRACTION LINE FROM HYD POWER PACK CHAFED THROUGH. HYDRAULIC LINE WAS CHAFING ON NOSE WHEEL WELL COVER INSIDE FWD BAGGAGE COMPARTMENT. REMOVED AND REPLACED LINE ASSY WITH LOCALLY MFG LINE ASSY AND REPOSITIONED TO PREVENT FURTHER CHAFING ON WHEEL WELL COVER. OPS AND LEAK CHECKED SATISFACTORY. (K)

IN LAST 6 TO 8 MONTHS, FLEET HAS HAD CRONIC PROBLEMS WITH TUBE FAILURES ON MAIN GEAR ONLY. INVESTIGATION SHOWS NO DAMAGE TO THE CORRESPONDING AREA OF THE TIRE INSIDE OR OUT. THE RUPTURE ON THE TUBES HAS BEEN VERY CONSISTENT. ALWAYS MAIN GEAR, ALWAYS O THE INBD SIDEOF THE TUBE, ALWAYS IN THE AREA WHERE THE TREAD OF THE TIRE TRANSITIONS INTO THE SIDEWALL AND ALWAYS A VERY SMALL SLIT OR PUNCTURE. HAVE EXAMINED INSTALLATION TECHNIQUES , INFLATION PRESSURES, AND LANDING TECHNIQUES AND DO NOT SUSPECT ANY OF THESE BEING THE CAUSE. HAVE BEEN IN CONTACT WITH MFG, THEY HAVE PROVIDED US WITH SOME TUBES MADE WITH A NEW COMPOUND WHICH HAVE FAILED ALSO. (K)

(CAN) INSP REVEALED THAT THE THROTTLE BRACKET WHICH IS RIVETED TO THE TOP OF THE CARB HEAT BOX WITH BLIND RIVETS HAD WORKED LOOSE AND WAS ALLOWING MIS-ALIGNMENT OF THE THROTTLE CONTROL CABLE AND CAUSING THE THROTTLE TO STICK.

AFTER A GROUND RUN UP, THE PILOT HAS NOTICED LARGE AMOUNTS OF OIL UNDER THE ACFT ON THE RAMP. FINDINGS WERE THAT THE OIL HAD COME OUT OF THE ENGINE BREATHER TUBE. UPON FURTHER INVESTIGATION, FOUND THAT THE BREATHER LINE WAS ROUTED SO THAT THERE WAS A "TRAP" IN THE LINE COLLECTING WATER AND FREEZING CAUSING THE CRANKCASE TO PRESSURIZE AND BLOW THE OIL OVERBOARD.

INSTRUCTOR NOTED ON A PREFLIGHT THAT THE LT AILERON WAS EMITTING A SQUEAKING SOUND . FOUND THAT IT WAS DUE TO A LACK OF LUBRICATION OF THE AILERON ROD END BEARING . NOTING THIS, DUE TO A EARLIER INSTANCE OF ANOTHER SN SAME TYPE ACFT THAT A STIFF AILERON CONTROL IN FLIGHT DUE TO SAME REASONS. LUBRICATION REMEDIED THIS PROBLEM.

AFTER A GROUND RUN UP, THE PILOT HAS NOTICED LARGE AMOUNTS OF OIL UNDER THE ACFT ON THE RAMP. FINDINGS WERE THAT THE OIL HAD COME OUT OF THE ENGINE BREATHER TUBE. UPON FURTHER INVESTIGATION, FOUND THAT THE BREATHER LINE WAS ROUTED SO THAT THERE WAS A "TRAP" IN THE LINE, COLLECTING WATER AND FREEZING CAUSING THE CRANKCASE TO PRESSURIZE AND BLOW THE OIL OVERBOARD.

DURING TAKEOFF ROLL, THE ENGINE ELECTRONIC CONTROL WENT OFF LINE AND THE CREW ABORTED THE TAKEOFF AFTER RETURNING TO THE RAMP, INSPECTION OF THE ENGINE SHOWED POWER TURBINE BLADE DAMAGE. THE ENGINE WILL BE REMOVED AND FORWARDED FOR INVESTIGATION AND REPAIR. P&WC WILL CONTINUE INVESTIGATING THE EVENT AND ADVISE OF ROOT CAUSE ONCE ESTABLISHED.

DURING A PHASE 1A INSPECTION, THE PROPELLER CONTROL BRACKET LOCATED UNDER THE CENTER PEDESTAL WAS FOUND TO HAVE 2 SHEARED OFF RIVETS. IN LIEU OF RIVETS IT WAS FOUND THAT 2 C-CLAMPS WERE HOLDING THIS BRACKET IN PLACE. WHAT IS IMPORTANT TO NOTE IS THAT THIS WAS THE FIRST PHASE INSPECTION DONE AT THIS FACILITY. PRIOR TO THIS THE AIRCRAFT UNDERWENT A FULL PHASE 1 TO 4 INSPECTION AT A BEECHCRAFT REPAIR STATION IMMEDIATELY BEFORE BEING IMPORTED. PLEASE SEE ATTACHED PICTURES AND DOCUMENTS.

AIRCRAFT WAS IN CRUISE AT FL210. FIRST OFFICER ATTEMPTED TO ADJUST THE RT POWER LEVER TO BRING THE TORQUE UP TO MATCH THE LT ENGINE. FOUND THE POWER LEVER ALREADY AT THE STOP. CREW THEN NOTICED A GRADUAL LOSS OF POWER (TORQUE, FUEL FLOW AND ITT) TO ABOUT 50 PERCENT TORQUE. THEY DEPLOYED THE ICE DOORS TO TROUBLESHOOT, AND AFTER SOME TIME, THE POWER RECOVERED TO 90 PERCENT. AFTER ABOUT 5 MINUTES, THE POWER BEGAN DECAYING AGAIN. AT 40 PERCENT TORQUE THE CAPTAIN INSTRUCTED THE F/O TO PREPARE FOR AN ENGINE FAILURE. AT APPROX 20 PERCENT TORQUE THE PROP AUTO-FEATHERED AND THE CREW SECURED THE ENGINE. AN UNEVENTFUL LANDING WAS MADE. MAINTENANCE TROUBLESHOT AND FOUND THE MAX FUEL FLOW AVAILABLE WAS APPROX 100 LB/HOUR, NOT ENOUGH TO EVEN GET THE ENGINE STARTED. THE FUEL CONTROL WAS REPLACED AND THE ENGINE GROUND RUN SERVICEABLE. PART IS BEING SENT TO STANDARD AERO FOR REPAIR. WILL ADVISE FURTHER DETAILS WHEN THEY BECOME AVAILABLE.

DUE TO PAST ISSUES WITH ELT`S GOING OFF, THE ELT`S WERE FOUND TO BE AT THE INCORRECT CLOCK ANGLE. THE ELT WAS REPOSITIONED TO THE PROPER ANGLE, WHICH WE THINK MAY HAVE CAUSED THE CRACKING OVER TIME. A SMALL CRACK WAS FOUND IN THE BOTTOM ANGLE BRACKET RADIATING FROM THE MIDDLE SCREW HOLE. THE ELT`S ARE NOW BEING REPLACED WITH THE KANNAD 406 ELT WHICH IS SUPPORTED DIFFERENTLY.

AUX FUEL PUMP FAILURE WITHIN 5.5 HOURS AFTER UNIT INSTALLED. THIS PUMP WAS ORIGINALLY REPAIRED FOR LOW PRESSURE OUTPUT AND HAS AGAIN FAILED TO PRODUCE PRESSURE. POSSIBLE QUALITY CONTROL ISSUE FROM OVERHAUL/REPAIR SHOP.

MAG WAS RECEIVED FOR AN OVERSPEED INSPECTION (500 HOUR INSPECTION PROCEDURE). BEFORE DISASSEMBLY THE BEARINGS WERE INSPECTED BY TURNING THE ROTATING MAGNET ASSEMBLY. SIDE PLAY WAS NOTED AND THE END PLAY WAS MEASURED AT 0.004 INCH. TCM DOES NOT ALLOW SIDE PLAY AND ONLY ALLOWS 0.0005 TO 0.0015 END PLAY. THE MAG WAS DISASSEMBLED AND THE BEARINGS WERE INSPECTED AND FOUND SERVICEABLE. THE 500 HOUR INSPECTION WAS CARRIED OUT WITH THE BEARINGS SHIMMED TO THE PROPER PRELOAD AND THE MAG WAS RETURNED TO SERVICE. THIS MAG WAS A FACTORY NEW UNIT WITH 412.7 HOURS IN SERVICE.

DURING A VISUAL INSPECTION ON A 100 HR, THE MAIN FUEL FEED LINE WAS FOUND SMOKING THOUGH THE AFT FIRE WALL. UPON FURTHER INSPECTION OF FUEL LINE THE SMALL GROMMET/BUSHING WAS FOUND WORN THROUGH CAUSING THE FIRE WALL SUPPORT TO CHAFE INTO THE LINE. THE LINE WAS REMOVED FROM SERVICE AND REPLACED NEW.

DUE TO PAST ISSUES WITH ELT`S GOING OFF, THE ELT`S WERE FOUND TO BE AT THE INCORRECT CLOCK ANGLE. THE ELT WAS REPOSITIONED TO THE PROPER ANGLE, WHICH WE THINK MAY HAVE CAUSED THE CRACK OVERTIME. A SMALL CRACK WAS FOUND IN THE BOTTOM ANGLE BRACKET RADIATING FROM THE MIDDLE SCREW HOLE. THE ELT`S ARE NOW BEING REPLACED WITH THE KANNAD 406 ELT WHICH IS SUPPORT DIFFERENTLY.

CORROSION WAS FOUND AT THE JUNCTION OF THE UPPER AND LOWER BODY OF THE FRONT SEAT CONTROL STICK ASSY. THE UPPER STICK HALF SLIDES INTO THE LOWER STICK HALF AND IS THEN HELD TOGETHER WITH A SINGLE THROUGH BOLT. THE CORROSION WAS DETECTED PROTRUDING FROM THE TOP OF THE JOINT WHERE THE UPPER PIECE SLIDES INTO THE LOWER PIECE OF THE ASSY. THE EXTENT OF THE CORROSION COULD NOT BE DETERMINED WITHOUT REMOVEING AND DISASSEMBLING THE STICK. CORROSION WAS SEVERE ENOUGH THAT SEPARATION OF THE (2) STICK PARTS WAS ACHIEVED ONLY AFTER APPLICATION OF HIGH HEAT FROM A TORCH. THE PROBABLE CAUSE IS PERSPIRATION FROM THE PILOT`S HAND DRIPPING DOWN THE STICK TO WHERE THE PARTS ARE MATED. CORROSION EASILY DEVELOPS WHEN THE MOISTURE CONTACTS THE BARE METAL IN THE JOINT. WHAT IS DISTRUBING IS THAT THE ACFT HAS ONLY BEEN IN SERVICE FOR LESS THAN 2 YEARS SINCE NEW. MFG SHOULD HAVE PROVIDED BETTER CORROSION PREVENTION OR DESIGN. SHOULD CORROSION BE ALLOWED TO CONTINUE, THAT THE CONTROL STICK WILL ULTIMATELY FAIL. INSP OF (2) OTHER ACFT OF THE SAME MAKE, MODEL AND IN-SERVICE TIME REVEALED SIMILAR CONDITIONS. (K)

PRIOR TO INSTALLING STARTER GEN SB 80.00.07 PARA 2B2B AND 2B3 WERE APPLIED. THE STARTER WAS THEN INSTALLED AS PER MM AND A GROUND RUN WAS DONE. DURING THE GROUND RUN NO VIBRATION WAS NOTED ON THE STARTER, VOLTAGE AND AMPERAGE WAS NORMAL. I NOTICED A BURNING ODOR SIMILAR TO WHEN THE ROTOR BRAKE WAS APPLIED AND SAW PARTS OF THE FRICTION RING OF THE DAMPER ASSY ON THE ENGINE DECK AGAIN NO ABNORMAL VIBRATION WAS NOTED ON THE STARTER. THE PILOT SHUT DOWN THE ENGINE. REINSTALLED THE OLD STARTER, STILL HAD TIME LEFT TO OH. UPON CLOSER INSPECTION IT WAS NOTED THAT THE FRICTION RING WAS COMPLETELY BROKEN UP. IT WAS SENT BACK TO APC FOR FURTHER INVESTIGATION.

DURING TAKEOFF (ASCENT), ACFT SUFFERED A BIRD STRIKE IN THE RT ENGINE. THE PILOT VERIFIED HIS PARAMETERS AND THEY DID NOT HAVE ANY VIBRATION. HOWEVER, HE SMELLED BURNED BIRD AND DECIDED TO RETURN TO DEPARTURE AIRPORT. DURING THE WALK AROUND INSP, THE PILOT EVIDENCE DAMAGES IN THE RT ENGINE FOR BIRD INGESTION (FIRST IMPELLER DAMAGE). THE ENGINE WAS TO BE SENT TO REPAIR STATION. (K)

IN CRUISE AT FL230 SOUTHBOUND, A LOUD BANG HEARD BY BOTH CREW AND SAW F/O SIDE WINDSHIELD SHATTER, BUT REMAINED INTACT. NO LOSS OF PRESSURIZATION. FLIGHT DESCENDED TO FL100 AND CREW INFORMED ATC BUT DID NOT DECLARE AN EMERGENCY. LANDED WITHOUT FURTHER INCIDENT. MAINTENANCE IS IN PROCESS OF REPLACING THE WINDSHIELD. BONDING CHECKS WERE PERFORMED ON THE WINDOW PRIOR TO REMOVAL AND WERE FOUND WITHIN ACCEPTABLE RESISTANCE LIMITS IN ACCORDANCE WITH M7 AEROSPACE SB 227-56-010 WHICH WAS ISSUED TO HELP ELIMINATE THIS TYPE OF FAILURE.

(CAN) DURING PREFLIGHT INSP ON NEWLY DELIVERED ACFT, DISCOVERED A .5 INCH COUNTERSINK RESTING UNDER ACFT TRANSMISSION. NO DAMAGE WAS FOUND BUT THE TOOL IS LARGE ENOUGH TO HAVE CAUSED FLIGHT CONTROL INTERFERENCE OR TO HAVE CONTACTED THE ROTOR BRAKE DISK. TOOL WAS REMOVED. NO FURTHER ACTION TAKEN.